czech aerospace - Výzkumný a zkušební letecký ústav

Transcription

Introductory Lecture to the ARC 2005
Slovo úvodem k semináři CLKV 2005
Application of Nelder-Mead Algorithm
in Aerodynamic Optimisation
Použití metody Nelder-Meada při
aerodynamické optimalizaci
FEM Based Fatigue Life Analysis
of Landing Gear
Analýza životnosti podvozku na
základě výpočtů MKP
Computer Simulation of Cabin
Environment in the Evektor EV-55 Aircraft
Composites — A New Trend in the
Airplane Control System
Compression Tests to Generate Materials
Property Data for Modelling FSW
Tandem Blade Centrifugal Compressor
Impeller Design
Počítačové modelování mikroklimatu
v kabině dopravního letadla
Evektor EV-55
Kompositní materiály — nový trend
v konstrukci řízení letounů
Tlakové zkoušky pro stanovení
materiálových charakteristik použitých
v numerickém modelu FSW
Návrh oběžného kola odstředivého
kompresoru s tandémovým
uspořádáním lopatek
Spectral Decomposition Use in Noise
Abatement of Propeller Driven Airplanes
Užití spektrální dekompozice při snižování
hluku vrtulových letadel
Introduction to Problems of Thermosetic
Composite Materials Recycling
Úvod do problematiky recyklace
termosetických kompositních materiálů
Innovation of MAC Microaccelerometer
Safety Analyses of Aircraft
Avionic Systems
Aerodynamic Design of V44 Model Propeller
and Aeroacoustic Characteristic Calculation
Inovace mikroakcelerometru MAC
Rozbor bezpečnosti avionických systémů
letadla
Aerodynamický návrh modelové vrtule V44
Výpočet aeroakustických charakteristik
N o v e m b e r
Contents / Obsah
2 0 0 5
ISSN 1211—877X
CZECH
AEROSPACE
Proceedings
LETECK Ý
zpravodaj
In this issue:
Czech Aerospace
Research Centre
CLKV
Proceedings of the
5th Annual
Workshop held at
Holany, Czech
Republic
November 3 to 4,
2005
Zatížení vodorovných ocasních ploch
při manévru
Centrum leteckého
a kosmického
výzkumu
Application of Artificial Neural Networks
for Gas Path Analysis of a Turbine Engine
Použití neuronových sítí při analýze
plynové cesty turbinového motoru
CLKV
Airfoil Pressure Distribution Measurement
on Ground Mobile Laboratory
Měření distribuce tlaku profilu na pozemní
pojízdné laboratoři
Load of Horizontal Tail by Manoeuvre
Analytical Methods for the Calculation of
Buckling in Composite Sandwich Panels
Výpočet vzpěrné pevnosti kompositových
sendvičových desek
Complex Equipment Certification from the
Reliability and Safety Points of View
Certifikace složitého vybavení z pohledu
zajištění spolehlivosti a bezpečnosti
© C Z E C H AE R O S PAC E
M A N U FAC T U R E R S A S S O C IAT I O N
Sborník vybraných
referátů
přednesených na
5. ročníku
semináře CLKV
Holany, 3. — 4.
listopadu 2005
No. 3 / 2005
Environment in Evektor EV-55 Aircraft
Colour illustrations to the article published on pages 9-10.
CZECH
AEROSPACE
Figure 4 — Temperature pattern, head
height, variety 1A
height, variety 1B
P r o c e e d i n g s
J OU R N A L
F O R
C Z E C H
AE RO S PAC E
R E S E A R C H
LETECK Ý
zpravodaj
VÝZKUMNÝ A ZKUŠEBNÍ LETECKÝ ÚSTAV, a.s.
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height, variety 2A
height, variety 2B
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Figure 8 — Airflow pattern, cross section, variety 1A
Figure 9 — Airflow pattern, cross section, variety 1B
Czech AEROSPACE Proceedings
Letecký zpravodaj
3/2005
© 2005 ALV / Association of Aviation Manufacturers, All rights reserved. No part of this publication may
be translated, reproduced, stored in a retrieval system or transmitted in any form or by any other means, electronic,
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ISSN 1211 - 877X
1
L E T E C K Ý Z P R AV O D A J
3/2005
Contents / Obsah
2
Aerospace Research Centre — Introductory Lecture at the 2005 ARC Workshop
Úvodní přednáška na Semináři CLKV 2005
Prof. Ing. Antonín Píštěk, CSc.
3
Application of Nelder-Mead Algorithm in Aerodynamic Optimization
Použití metody Nelder-Meada při aerodynamické optimalizaci
Ing. Zbyněk Hrnčíř, Ing. Daniel Langr, Ing. Robert Popela, PhD
6
FEM Based Fatigue Life Analysis of Landing Gear
Analýza životnosti podvozku na základě výpočtů MKP
Ing. Petr Augustin, Ph.D., Ing. Martin Plhal, Ph.D., Ing. Jan Šplíchal
9
Computer Simulation of Cabin Environment in the Evektor EV-55 Aircraft
Počítačové modelování mikroklimatu v kabině dopravního letadla Evektor EV-55
Ing. Jan Fišer
11
Composites — A New Trend in the Airplane Control System
Kompositní materiály — Nový trend v konstrukci řízení letounů
Ing. Tomáš Marczi, MSc, PhD, Ing. Miroslav Kábrt
14
Compression Tests to Generate Materials Property Data for Modelling FSW
Tlakové zkoušky pro stanovení materiálových charakteristik použitých v numerickém modelu FSW
Ing. Petr Bělský
17
Tandem Blade Centrifugal Compressor Impeller Design
Návrh oběžného kola odstředivého kompresoru s tandémovým uspořádáním lopatek
Ing. Daniel Hanus, CSc, Ing. Tomáš Čenský, Ing. Ivan Jeřábek
22
Spectral Decomposition Use in Noise Abatement of Propeller Driven Airplanes
Užití spektrální dekompozice při snižování hluku vrtulových letadel
Ing. Tomáš Salava, DrSc, Ing. Marcela Šloufová
26
Introduction to Problems of Thermosetic Composite Materials Recycling
Úvod do problematiky recyklace termosetických kompozitních materiálů
Mroslav Valeš, Ing. Petr Kachlík
28
Ing. Milan Chvojka, Ing. Josef Fabián
31
Safety Analyses of Aircraft Avionic Systems
Rozbor bezpečnosti avionických systémů letadla
Ing. Jiří Hlinka, PhD
34
Aerodynamic Design of V44 Model Propeller and Aeroacoustic Characteristics Calculation
Aerodynamický návrh modelové vrtule V44. Výpočet aeroakustických charakteristik
Ing. Jan Dostál, Ing. Pavel Klínek
40
Zatížení vodorovných ocasních ploch při manévru
Ing. Ivo Jebáček
41
Application of Artificial Neural Networks for Gas Path Analysis of a Turbine Engine
Použití neuronových sítí při analýze plynové cesty turbinového motoru
Ing. Jaromír Lamka
44
Airfoil Pressure Distribution Measurement on Ground Mobile Laboratory
Měření distribuce tlaku profilu na pozemní pojízdné laboratoři
Ing. Jan Friedl, Ing. Martin Kouřil, Ph.D., Ing. Róbert Šošovička, Ph.D.
46
Analytical Methods for the Calculation of Buckling in Composite Sandwich Panels
Výpočet vzpěrné pevnosti kompozitových sendvičových desek
Ing. Martin Baumruk
51
Complex Equipment Certification from the Reliability and Safety Assurance Points of View
Certifikace složitého vybavení z pohledu zajištění spolehlivosti a bezpečnosti
Ing. Milan Merkl, CSc; Miroslava Nová
2
C Z E C H A E R O S PA C E P R O C E E D I N G S
Aerospace Research
Centre
Introductory Lecture at the 2005 ARC
Workshop
Prof. Ing. Antonín Píštěk, CSc, Director, Institute of Aerospace Engineering,
Brno University of Technology
Ladies and Gentlemen, Dear Colleagues,
We are opening this years’s edition of traditional Aerospace Research
Centre workshops to present results achieved in 2005, the first year of
the existence of a new Centre. When I opened the 2004 workshop we
were just before the final evaluation of its activities and were awaiting
a decision on starting a new Centre. The evaluation was very successful, with the Centre getting maximum ”points“ and being placed in the
first position among other research centres candidates. So the necessary requirements have been met for the new Centre to continue work
in 2005 till 2009.
The objectives of the new Centre have not changed much, its mission being again the concentration of research capacities and technology transfer of research results to the subjects that will use them in
practice. The work done by the previous Centre has already proved
the beneficial composition of research teams consisting of specialists
from universities, research institutes and aerospace industries. So it
was logical to preserve this well-established organizational structure.
One of the appreciated benefits of the previous Centre was participation of young specialists in particular projects solved on a broader
basis of European research programmes. There were some drawbacks
too, often commented on at ARC workshops. One of them was socalled ”series structure“ when tackling particular tasks, which was
difficult to coordinate and which often led to dualities at work. For this
reason an essential change has been made. The original division of
projects into particular tasks B1 through B5 assigned to VUT Brno
(Brno Technical University), V1 through V13 assigned to VZLÚ
Praha (Aeronautical Research and Test Institute, Prague), and P1 and
P2 for ČVUT Praha (Prague Technical University) was replaced by
a matrix block structure A, B, C, D, E, F on the solution of which all
the organizations will participate with certain priorities. This workshop
already reflects this fact in that authors of some papers are from different workplaces.
The matrix structure can be seen in the following diagram.
stries’ representatives in the ARC Council, who now have more say in
influencing research programmes and can help in economic agenda
and staffing. Tasks of the aviation industry may be summarized as follows:
— Defining requirements for application research and project feasibility
— Practical application of research results
— Training of ARC researchers (graduate studies)
— Use of labs and software for education of workers from industry,
etc.
Nowadays ARC is working on several projects of the aviation industry. They include particularly the VUT 100 airplane, now under certification, a new airplane EV-55, which is seen as a project of the Czech
aviation industry under the umbrella of the Association of Aviation
Manufacturers, a new all-composite glider G 304 S, and a number of
other smaller projects.
The matrix structure of tasks division is sophisticated in terms of
work coordination on individual blocks. That is why a coordination
team has been set up made up of young and skilled representatives of
ARC. The today’s workshop is managed by this team.
It is mainly young researchers that are doing research for ARC.
Their education courses are part of ARC research programme.
Current status and structure of ARC employees:
68 employees
38.1 average age
Age diagram of ARC employee
7 employee
ow er 60 years
8 employee
51-60 yaers
Participation in
European
structures
up to 20
6 employee
41-50 years
29 employee
21-30 years
21-30
31-40
41-50
Nowadays the process
51-60
60 and more
18 employee
of project evaluation is
31-40 years
coming to an end, The
projects in question
were submitted in the
last call of EU 6th Framework Programme, and for the Ministry of Industry and Trade
grants. Let us leave judging the results for the future. What is clear is
that thanks to active participation of research organizations in various
European committees and bodies the researches are often invited to
take part in the EU projects. The first large Czech project coordinated
by VZLU has a chance to be adopted, which would give ARC the
opportunity to take part in working on it.
Conclusion
Another improvement on the previous managerial structure is a better
link with the aerospace industry through direct engagement of indu-
I cannot help repeating once more that aerospace research has a big
impact on technology development. This is evidenced by huge EU
funds that go to research and development in aeronautics and astronautics. The existence of ARC is one of the guarantees of our participation in these efforts.
3
3/2005
in Aerodynamic Optimization
Použití metody Nelder-Meada při aerodynamické optimalizaci
Ing. Zbyněk Hrnčíř, Ing. Daniel Langr, VZLU, Plc., Prague; Ing. Robert
Popela, PhD, Brno University of Technology
An aerodynamic optimization tool, based on Nelder-Mead simplex method has been developed. The software couples the optimization algorithm with three CFD solvers and three grid generators used at VZLU. After various validation test cases the tool was
successfully applied to optimize transonic wing planform within the EU project VELA (Very Efficient Large Aircraft). The lift to
drag ratio of outer wing was improved by 38% by application of the method.
Byl vyvinut nástroj pro aerodynamickou optimalizaci založený na simplexové metodě Nelder-Meada. Tento program spolupracuje
s třemi CFD výpočetními programy a třemi generátory síti v současné době používanými ve VZLU. Po řadě testovacích úloh byl
tento program využit při aerodynamické optimalizaci transsonického křídla v rámci evropského projektu VELA (Very Efficient
Large Aircraft). Aplikací prezentované metody se podařilo zvýšit aerodynamickou účinnost vnější části křídla o 38%.
Keywords: CFD, aerodynamics, optimization, simplex method.
Introduction
Aerodynamic optimization methods are crucial tools in the aircraft
development process. In today's competitive environment it is not
possible to rely on traditional trial an error method even if it is done
by experienced aerodynamicists. Uncompromising demands on performance parameters and design cycle time require a fully automatic
tools which would allow the designer to explore design space and
efficiently reach the optimal solution. Nowadays, when Computational Fluid Dynamics (CFD) matured point, where it is routinely used
for aerodynamic analysis of future aircrafts and calculation cost is falling as computational power is rapidly growing, simulation-based
optimization is more and more promising and accessible. Lots of optimization approaches are available today in the branch of aerodynamic
design. They extend from direct search methods, through gradient
based methods and genetic algorithms to quite new approaches like
automatic differentiation or control theory methods. The selection of
the most appropriate method highly depends on properties of a specific problem we want to solve. The most important ones are: computational cost of fitness evaluation, number of design parameters, and
presumptive shape of fitness function. The application of NelderMead simplex method (NMSM) which belongs to the category of
direct search approaches will be demonstrated in the following.
Nelder-Mead simplex method
The Nelder-Mead simplex method was first published in 1965 by
J. A. Nelder and R. Mead [1]. It is a direct local descent algorithm,
which does not use the fitness function derivatives and the fitness
function does not need to be smooth. On the other hand it tends to be
efficient only for function of relatively low dimensions. Acceptable
number of dimensions is up to 10. A ”simplex“ is a geometrical figure formed by n+1points in n-dimensional design space. For example,
if we want to find the best position of flap relative to main element,
we work with two design parameters, the horizontal and the vertical
location of flap. It means, the design space is two-dimensional and
simplex forms a triangle. If we decide to add next variable such as
angle of flap deflection, the simplex would be a tetrahedron. At the
beginning of the searching process, the initial simplex is randomly
generated and the fitness function is evaluated at all of its vertexes.
Through a sequence of geometric transformations such as reflection,
contraction, expansion and multi-contraction is the initial simplex
modified and the vertex with the worse fitness value is replaced by
new better one. At each step it is necessary to check that the genera-
ted new point is inside limits of design space. This method is usable
only for finding local minima of functions. Due to the continual change of simplex shape during search process the method is also known
as Amoeba, referring to a large protozoan (Amoeba proteus) who continually changes it's shape.
Description of developed optimization tool
The presented optimization tool is based on NMSM and it interconnects several CFD solvers and grid generators. It allows user to choose suitable solver or grid generator according to the problem which is
to be solved. At this time the optimization algorithm can be used with
three CFD solvers: Edge from FOI for RANS or Euler calculations on
general geometries and grids, Fluent from Fluent Inc. for similar
application as pervious one and MSES written by Prof. M. Drela of
MIT for 2D analysis of single or multi-element airfoils. As a grid
generators it is possible to use TGrid from Fluent Inc. for 2D triangular grids, ICEM Tetra and ICEM HEXA both from ANSYS for tetrahedral and hexahedral grids. The practical application of the tool will
be demonstrated on transonic wing planform optimization.
Transonic wing planform optimization
VZLU was involved in 5 FP EU project VELA (Very Efficient Large
Aircraft), which was finished successfully in September 2005. This
project was focused on exploration of flying wing configuration from
multidisciplinary point of view. It is believed that it is a promising and
possible configuration for the future aircraft. In this project, we were
responsible for aerodynamic optimization of the unconventional aircraft at cruise speed Mach = 0.9 at operational altitude 35,000 ft. The
optimization presented here was focused only on outer part of the flying wing planform and served as a testing case before more complex
optimization of whole VELA configuration which was involving
moreover 2d optimization of wing sections using hybrid genetic algorithm and inner part of wing body was taken into account as well.
Fig. 1. Initial VELA flying-wing configuration
4
Fig. 2. Outer part of wing with six design variables
Geometry and parameterization
The geometry of outer wing was composed of two linear surfaces
which were formed by three airfoils. The chord lengths and twist angles of the three sections were considered as design variables and were
subjects of our optimization. There were a few geometrical constraints.
1. the span position of the three basic airfoils was constant
2. the sweep angle of leading edge was 45° and was constant
3. the chord of second airfoil (variable Chord2) must be in interval
<Chord1,Chord3>
4. the chord lengths must be smaller than 150% and longer than
50% of original length
5. twist angle is done about 25% axis and angles must be inside the
range of original twist angle value ± 5°
Fitness function and its evaluation
The aim of the effort was to reshape the wing planform in the way
which would ensure the highest possible aerodynamic efficiency at
cruising conditions. The initial wing was working at lift coefficient of
0.33 which is related to certain lift force L. During the optimization
process, the wing area (A) is not constant due to the changes of chord
lengths. It means, the aerodynamic efficiency of proposed wing must
be evaluated at angle of attack (AoA) related to the lift force L of the
initial wing. In other words, the AoA is not constant during optimization process and must be found the right one for each proposed wing
planform. It ensures that proposed wings are producing the same lift
force as initial wing and thus they are comparable. The searching of
the proper AoA was done through the following iteration process. At
first, the optimization algorithm proposed all six design variables for
new wing shape. According to the area of the new wing the desired
operating CLoper was determined.
When the AoA corresponding to the desired lift force was found,
it was possible to evaluate the aerodynamic efficiency. For this purpose the wing drag needed to be determined. Due to the large amount of
computations needed to be performed during optimization process we
decided to use flow solver in euler mode, it means no viscous effects
were taken into account. The solution time of euler calculations is
considerably shorter in comparison with viscous (RANS) calculation.
From physical point of view, it is possible to justify this decision by
very high Reynolds number (Re=7.107) and by supposed small
amount of separated flow on such stream-line body. Since the drag
resulting from euler solution does not contain skin friction drag, it is
necessary to compute this component separately. To estimate friction
drag of the wing, viscous drag of turbulent flat plate for both wing
sections is derived and summed using following formulas:
Average friction coefficient on turbulent smooth flat plate for high Re
(Re > 107), [2]
Cfav = 0,455 / (log Re)2,58
and friction drag for entire wing,
CDf = 2. (cfav1.A1 + cfav2.A2) / (A1 + A2)
This leads to a friction drag of about 30 % of total drag. The estimated friction drag is added to the drag calculated by euler solver (contribution of induced and wave drag). The fitness function is then simple lift to drag ratio representing aerodynamic efficiency multiplied by
-1, because the optimization algorithm is set to searching a minimum
not a maximum.
F = -1. cL / cD
Since the flow field is transonic, the presence of shock waves could be
expected. The interaction of shock waves with boundary layer can
cause the massive flow separation and this situation can lead to substantial performance deterioration which can not be predicted by inviscid euler solver. Due to this danger it was necessary to check the final
design by viscous RANS solver, where the viscous and turbulence
effects were included.
Computational grid
ICEM CFD software was used for parametric generation of computational grid. The grid is automatically generated according to a script
CLoper = L / (1/2. ρ. v2 . Anew) = CLorig . Aorig / Anew
Then the desired AoA was found in three steps as depicts Figure 3.
Fig. 3.Iterative approach for searching the AoA relating to desired
operating CL
Fig. 4. Grid on wing for euler (top) and RANS (below) calculation.
5
file. In the script file the values for all the six design variables are defined. When optimization algorithm changes the values in script file,
ICEM will generate new grid corresponding to the new wing shape
without any user's action. There is no doubt that the grid quality is fundamental parameter affecting the accuracy of calculations, especially
optimization processes are very sensitive to accuracy and level of convergence. For these reasons we decided to use hexahedral grids. In
case of such simple and slender geometry, it is a logical choice. By
using hexahedral structure it is possible to effectively distribute grid
density, the fine resolution in chord-wise direction to catch properly
sock wave features and relatively coarse grid distribution in span-wise
direction to reduce number of cells as much as possible.
3/2005
[CL/CD]
Number of fitness evaluations
Flow solver
As a flow solver was used EDGE developed at FOI. It is an unstructured flow solver for arbitrary elements. The flow solver uses an
edge-based node-centred finite volume method to solve the governing
equations. The control volumes are non-overlapping and are formed
by a dual grid, which is computed from the control surfaces for each
edge of the primary input mesh. The governing equations are integrated explicitly towards steady state with multi stage Runge-Kutta time
integration. Convergence is accelerated by agglomeration multigrid
and an implicit residual smoothing. During the optimization process
the solver was used in euler mode. For the performance check of the
final design the RANS mode was used. In this case the one - equation Spalart-Allmaras turbulence model was employed.
Fig. 6. Development of fitness function (lift to drag ratio). Inviscid
results.
Optimization run
301
0.33
3°
17.52
The development of design variables during optimization process is
depicted in Fig. 5. The development of fitness function, the lift to drag
ratio, is depicted in Fig. 6. It is possible to see that the convergence of
the fitness function was reached after circa 90 steps, but each step
consists of three flow analyses of proposed wing at different AoA as
M=0,9;RANS
Area [m2]
274
CLoper
0.36
AoA
3.5°
CL/CD
24.25
Fig. 7. Comparison of the initial wing (left) and the final wing proposed
by optimization (right). Upper surface (top) and lower surface (below)
was mentioned afore. It means the optimization run performed 270
CFD analyses. It took about 80 hours on Dell power edge server with
four processors Itanium2 1.5 Ghz.
[m]
Conclusions
[°]
An aerodynamic optimization tool based on Nelder-Mead algorithm
was developed. The applicability of the tool was demonstrated on
aerodynamic optimization of transonic wing planform. The aim of the
optimization was to find the combination of chord lengths and twist
angles corresponding to the highest lift to drag ratio. The ratio was
improved by 38%. It should be noted that only aerodynamic point of
view was taken into account, neither structural nor handling characteristics. The optimization method should be used for searching minimum of low-dimensional functions with one global minimum. The
recommended number of dimensions is up to 10. If it is not known
how many local extremes the function has, the method can be used,
but it should be run several times from different initial simplex. Even
then there is no certainty that the method will converge to global
minimum.
References:
Fig.5. Development of design variables, chord lengths (top), twist angles (below).
[1]
Nelder J.A., Mead R.; A simplex method for function minimization; The Computer Journal 7 (1965) p.308-313
[2]
White, F.M., Fluid Mechanics; McGraw Hill, 4th. edition,
1999
6
of Landing Gear
Analýza životnosti podvozku na základě výpočtů MKP
Ing. Petr Augustin, Ph.D., Ing. Martin Plhal, Ph.D., Ing. Jan Šplíchal,
Institute of Aerospace Engineering, Brno University of Technology
The paper deals with the fatigue evaluation of nose and main landing gear of the VUT 100 Cobra airplane based on results of the
FEM stress analysis. The main goal is to determine the stress levels ensuring the requested undercarriage safe life and establish the
preliminary life value before starting the fatigue tests. A realistic random sequence of flights and loads within is applied in the fatigue
analysis considering essential service load conditions. The fatigue calculation itself is realized by the FEFAT module of the
MSC.Fatigue software. The fatigue analysis has provided the first evidence of possibility to reach the requested safe life, an overview
of critical locations of structure and recommendations for design improvements.
Příspěvek se zabývá výpočtem životnosti příďového a hlavního podvozku letounu VUT100 založeným na výsledcích napěťové analýzy
provedené pomocí MKP. Cílem je určení hladin napětí, které umožní dosažení požadované bezpečné životnosti a stanovení
předběžné životnosti do doby, než budou uskutečněny únavové zkoušky. Pro účely únavového výpočtu je uvažováno zatěžování
podvozků sekvencí letů s náhodným pořadím zatížení, která zahrnuje podstatné provozní zatěžovací případy. Vlastní výpočet je
prováděn s využitím modulu FEFAT programu MSC.Fatigue. Uskutečněná analýza prokázala možnost dosažení požadované
bezpečné životnosti podvozků, poskytla přehled o kritických místech a doporučení pro konstrukční modifikace.
Keywords: landing gear, fatigue, life analysis, FEM, MSC.Fatigue, FEFAT.
Introduction
During the current development of main and nose landing gear of
a new airplane VUT100 Cobra, a fatigue analysis has been started in
order to determine acceptable stress levels ensuring the requested
undercarriage life of 20 000 landings. Another aim of this evaluation
is to approve the preliminary value of safe life before starting the fatigue tests. This work is carried out at Institute of Aerospace Engineering in collaboration with EVEKTOR Ltd and Technometra Radotin
Plc.
The solution of this problem is based on finite element models of
landing gear prepared previously for the static analysis [1]. The fatigue calculation is realized by the FEFAT module of the MSC.Fatige
software [2].
Definition of load spectra
The VUT100 Cobra is designed for purposes ranged from basic and
advanced training of private and military pilots, touring and sport flying to general commercial use. With respect to certification of the airplane according to JAR23 / FAR23 standards, the ground load spect-
ra recommended by FAA for this type of airplanes are applied [3, 4].
The taxi and landing impact spectra are shown in Figure 1. Load cases
considered in the analysis involve main service conditions represented
by two types of landing impact, rebound, taxiing and turning [5]. For
purposes of the fatigue calculation, a realistic load history considering
defined spectra and multiple load cases was generated using the inhouse software FLTSIM [6]. This approach along with the application of the rain flow counting procedure allows correct analysis of
a multiple- load-case problem with realistic involvement of groundair-ground cycles that in some locations of structure represent an
essential portion of fatigue damage. The load sequence consists of
thousands of different flights and selection of loads is done on a random basis. Figure 2 presents the load history for fatigue analysis.
Besides the input files of the MSC.Fatigue software containing
load histories pertaining to x, y and z axes, the FLTSIM program was
also used for generation of nominal stress histories based on nominal
stress magnitudes defined for the load cases applied. It has allowed us
to perform an additional common fatigue evaluation of lugs using the
component S-N curve by Heywood. Another product of the FLTSIM
program is an input file for the control software of the servohydraulic
testing machine. With respect to the fact that the sequence should be
practically applicable to a fatigue test, the small amplitude taxi loads
were omitted and related fatigue damage was realized on one higher
load level using essentially smaller number of cycles. It will allow
acceptable time of fatigue test but also reduces computing time in the
case of detailed FEM model of the landing gear.
Description of FEM models
FEM model of main landing gear is presented in Figure 3. It consists
of ten main parts connected with respect to the contact. 398,000 quadratic volume elements were used in entire model. In the case of nose
landing gear shown in Figure 4, 394,000 various types of elements
were used and the model consists of fifteen main parts.
Calculation of stress-time histories
Fig. 1 — Ground load spectra
As the first step in the simulation of real loading of undercarriage, the
maximum principal stresses related to unit loads defined in three
7
3/2005
Fig. 3 — FEM model of
main landing gear
Fig. 2 — Samples of load-time histories applied to the FEM model of
nose landing gear
directions were calculated by MSC.Nastran using linear static analysis. MSC.Fatigue utilizes the principle of linear superposition to combine all load cases together and determine the stress variation in each
node of the FEM model (see Figure 5). This can be expressed by formula:
(1)
where the elastic stresses σij from each load case k are normalized by
the load magnitude from the FE analysis Fk and then multiplied by the
Fig. 4 — FEM model of
nose landing gear
time variation of the loading Fk(t). Obtained stress-time histories are
subsequently treated using the rain flow counting procedure and
extracted cycles are written into the rain flow matrix.
Calculation of fatigue damage
The MSC.Fatigue module allows the safe-life calculation based either
on the stress-life (S-N) or the strain-life (ε-N) fatigue curves. The ε-N
crack initiation approach was in this case hardly practicable because
Fig. 5 — Stress-time history in different locations of main landing gear
FEM model
α = 1)
Fig. 6 — Material S-N curves (α
8
Fig. 9 — Example of single location damage histogram
Fig. 7 — Log of damage contour plot for the cylinder of main landing
gear
curves of equivalent steels AISI 4130 and 14331 respectively had to
be used. The S-N curves are presented in Figure 6.
The calculation of fatigue damage is performed after a mean stress
correction using rain flow matrix occurrences and fatigue curves.
Figures 7 through 9 depict resulting damage maps and an example of
the single location damage histogram. Subsequent determination of
the mean fatigue life utilizes the Palmgren-Miner linear damage
accumulation hypothesis.
Conclusion
The fatigue evaluation of main and nose landing gear of the VUT100
airplane has been performed in order to determine the acceptable
stress levels and the preliminary safe life. The solution is based on
FEM models prepared for the static analysis and the stress-life approach supported by the MSC.Fatigue software. The fatigue analysis
has provided the first evidence of possibility to reach the requested
lifetime, an overview of critical locations of structure and recommendations for design improvements.
References:
Fig. 8 — Log of damage contour plot for the nose landing gear
console
of the lack of necessary material data. Besides the strain-life diagrams, the cyclic stress-strain curves have also to be defined before
a calculation is started. However, infrequent examples of using this
approach for the safe life evaluation of aircraft structures can be
found. The reasons mentioned above led to the application of S-N curves and the local elastic stress approach. The local elastic FE stresses
at the notches are directly applied for determining the number of cycles from the material S-N curve related to the stress concentration factor α of 1. Thus, MSC.Fatigue uses rather conservative technique
since although the linear FE stresses can be perceived as multiples
of nominal stresses, the fatigue curve of notched specimen is derived
from the unnotched one by division of stress amplitudes by fatigue
notch factor β(N), whereas in general β(N) < α.
Primary structure of landing gear is made of L-ROL and L-CM3
steels and 2024-T3 aluminum alloy. Suitable S-N curves were found
in references [7] and [8]. In the case of L-CM3 and L-ROL, the S-N
[1]
[1] Plhal, M.: Static Analysis of VUT100 Aircraft Main and
Nose-wheel Landing Gear; in Letecky zpravodaj, No.
3/2003, pp. 9-11.
[2]
MSC.Fatigue User's Guide; MSC.Software Corp.
[3]
Fatigue Evaluation of Wing and Associated Structure on
Small Airplanes, AFS-120-73-2, FAA, 1973.
[4]
Belohradsky, T.: Data for Safe Life Calculation and Fatigue
Tests of VUT100 Airplane Landing Gear; Report
EVUT991.10-ST, Evektor Ltd (in Czech).
[5]
Vychopen, J., Belohradsky, T.: Load Cases for Safe Life
Calculation and Fatigue Tests of VUT100 Airplane Landing
Gear; Report EVUT019.05-ST, Evektor Ltd (in Czech).
[6]
Augustin, P.: Flight by Flight Fatigue Test with Random
Application of Loads; in Letecky zpravodaj, No. 3/2003, pp.
3-6.
[7]
MIL-HDBK-5H Military Handbook, Metallic Materials and
Elements for Aerospace Vehicle Structures.
[8]
Becvarik, P.: Fatigue Properties of Czechoslovak Steels;
SNTL Prague, 1985 (in Czech).
9
3/2005
Počítačové modelování mikroklimatu v kabině dopravního
letadla Evektor EV-55
Ing. Jan Fišer / Institute of Aerospace Engineering, Brno University of
Technology
This paper deals with CFD simulation of chosen environment parameters in Evektor EV-55 aircraft cabin.
Simulations were focused on getting detailed information about temperature and airflow patterns which will be
created by air-condition system and also on verifying necessary heating output. A model was created using Star-CD
ver. 3.22 (pre/postprocessor, solver) software package.
Tento příspěvek se zabývá CFD simulací vybraných parametrů mikroklimatu v kabině připravovaného dopravního letadla Evektor
EV-55. Simulace byly zaměřeny na získání konkrétnější představy o teplotních a proudových polích, které bude navržený
klimatizační systém v kabině letadla vytvářet a také na ověření nezbytného tepelného výkonu pro vytápění kabiny. Modelování a
simulace byly provedeny pomocí programu Star-CD ver. 3.22.
Keywords: Cabin environment, CFD simulation, ECS, air-condition.
Introduction
Commercial transport aircraft operate in a physical environment that
is not survivable by unprotected humans. This requires a complex
ECS (environmental control system) to provide passengers and crew
with safety and comfort without health risk [1]. The ECS must also
create optimal airflow and temperature patterns to provide thermal
comfort.
Environmental control system in EV-55
The cabin is not pressurized, so ambient air can be used for ventilation and heating without compression. Air comes to the system
through inlet (5 — see Fig.1), which is equipped with a flow control valve and drain. Sufficient air amount during ground operation
is provided by a fan which is located in bypass. Hot bleed air is
mixed with ambient fresh air in the mix manifold (6). The mixed air
temperature is controlled by a bleed air regulating valve (4). Its
position is controlled by a temperature regulator, which evaluates
air temperature in the cabin, air temperature behind the noise suppressor (7) and the position of the temperature regulating button.
Then air passes through a distribution valve (9). Its position sets the
air amount proportion between the passenger cabin and the cockpit.
Air is supplied to the passenger cabin from nine distribution outlets
(8) located below seats. Four distribution outlets are located in the
cockpit; two are located near the floor (10) and two below the front
window (11). The ventilation system also contains nine individual
overhead air showers in the passenger cabin and two in the cockpit.
An ambient air enters the shower system by NACA inlets located
on fuselage (see Fig. 2). Air is exhausted form the whole cabin by
one outflow vent only, which is located in aft cargo compartment.
Aircraft can be equipped with a cooling air-condition unit. Cooling
heat exchangers are located in ceiling, compressor unit in aft cargo
compartment. The system works on circulating principle; warm air
is sucked up from the cabin, cooled-down in heat exchangers and
returned to the cabin.
Model description and model varieties
A CFD model of the passenger cabin was created using Star-CD 3.22
software package. Dimensions of model geometry are based on real
cabin dimensions. The whole volume of the geometry model was subdivided into a grid of cells (see Fig. 3). Some parts of the grid were
divided more, because higher gradients of temperature and velocities
were supposed there. This gave better results in these parts and also
supported convergence of solution. Final CFD model contains c.
150 000 cells.
Fig. 1 — EV-55 ECS (main parts)
Four model varieties (see Tab. 1) were chosen after consultation
with EVEKTOR ltd. staff. The first variety (label 1A) is for the
ground operation and extreme design climate conditions. Temperatu-
Tab. 1 — Model varieties
10
Tab. 2 - Simulations results
re and altitude are taken from the aircraft operation envelope. Variety 1B is for ground operation and standard design
climate conditions. The temperature is taken from international standard atmosphere. Varieties 2A and 2B are proposed
for flight operation. Variety 2A is for extreme climate conditions in typical cruise altitude; variety 2B is for standard
climate conditions. In all model varieties heating is required.
Air showers were not modelled.
Boundary conditions
The right specification of appropriate boundary conditions is
a very important part of every CFD simulation. The main
goal in this case was the right specification of thermal resistances, necessary heat output for heating and volume of
supplied air. Outlet velocity from distribution outlets was
Fig. 2 - Location of ECS parts
— The significant lengthwise airflow
rises in the cabin, because air is
exhausted by one outflow vent only.
Due to the lengthwise airflow, temperature in aft section of the cabin is
higher than the temperature in fore
section.
— The results will be used to change
some design details of the air-condition system and air distribution outlets
Fig. 3 - CFD model of passenger cabin
calculated from supplied air volume (2500 dm3/min) and total area of
outlets. The temperature of supplied air was calculated from heat loss.
K-Epsilon low Reynolds number turbulence model was used for turbulence modelling.
Results
Temperature patterns in head height are in Figures 4 to 7 and airflow
patterns in cross section (first row) are in Figures 8 to 11 (Editorial
note: The afore mentioned illustrations are printed in colour for better clarity on inner side of back cover). Average values of chosen environmental parameters are in Table 2 (see Tab. 2).
Conclusion
— The necessary heat output range of the air-condition system in the
heating mode is from 270 to 2600 W.
References:
[1]
Space, D. R., Waters, K. L., Willes, D. L.: HVAC applications; ASHRAE, Atlanta 1999
[2]
Carlie, D.: Cabin air comfort; FAST 19, Airbus 1996
[3]
Versteeg, H. K., Malalasekera, W.: Computation fluid dynamics; Peterson Education, Harlow 1995
[4]
Myška, J.: Metodika stanovení mikroklimatických podmínek
klimatizovaného prostoru letadla, (Establishing cabin
atmosphere in a nonpressurized aircraft); Sborník konference klimatizace a větrání, Společnost pro techniku prostředí,
Praha 2004 (in Czech)
11
3/2005
Composites — A New Trend in the Airplane
Control System
Kompositní materiály — Nový trend v konstrukci řízení letounů
Ing. Tomáš Marczi, MSc Ph.D. / Czech Technical University in Prague,
Department of Automotive and Aerospace Engineering, Division of
Aerospace; Ing. Miroslav Kábrt / Vanessa Air Ltd. Litomyšl
Main idea of the presented project is to develop and investigate mechanical properties of the airplane control system
composite connecting rod. The project is focused on the buckling stability of composite connecting rod and strength of
the connection of composite tube with metal end joint. Both project objectives will be investigated practically by
performed tests. The buckling stability of the composite tube depends on the filament, filament layout and used resin.
The strength of the connection of the metal end joint with the composite tube depends on the end joint geometry and
type of connection
Hlavním cílem projektu je vývoj a výzkum mechanických vlastností kompositní konstrukce táhel řízení letounu. Projekt je zaměřen
na vyšetřování vzpěrné stability kompositových táhel, pevnostních únavových a spolehlivostních charakteristik spoje: kompositová
trubka + kovová koncovka. Vyšetřování mechanických vlastností kompositových táhel se předpokládá pomocí praktických zkoušek.
Vzpěrná pevnost kompositového táhla závisí na druhu a materiálu použitých vláken, uspořádání vláken a na použité pryskyřici.
Pevnost spojení kovové koncovky s kompositovou trubkou závisí na geometrii a typu spoje.
Keywords: Control Rod, Composite, Carbon, Glass, Filament, Airplane Control System, Winding,
Gluing.
Introduction
Aircraft Gross Weight
Successful airplane control systems used so far are based on the system of control ropes or rods. Control system built from metal ropes is
called 'flexible', system build from connecting rods is known as 'rigid'.
Both of these systems have their advantages and disadvantages. The
'rigid' control system does not elongate during the service life, has
lower friction, is more durable than 'flexible' control system. On the
other hand, 'rigid' system is heavier and more time-consuming during
the manufacturing process.
Since the last decades of the previous century, the airplane industry has been much influenced by the composite materials. Their ratio
stress vs. weight vs. work difficulty vs. price is unbeatable by common
metal materials. At present, composite materials are replacing the
conventional materials in all parts of airplane structure. The aim of
this project is to develop and verify reliability of the composite connecting rod and its suitability for airplanes control systems, or any
other applications. The project is mainly focused on the problems
related to the composite connecting rod stability (buckling) and on the
strength of the connection: composite tube vs. metal head lug.
The common knowledge is that aircraft weight is one of the most
important parameters. By the late 1930s most of the major aircraft
companies had separate weight engineering staff. If the weight of
some structure component is increased for some reason and consequently it means adding weight elsewhere, it consequently leads to increase of airplane gross weight (Marczi T., 2002).
As already mentioned, the 'flexible' control systems are less effective and suitable for today's sport or GA airplanes. Most planes from
these categories use the 'rigid' control systems. The 'rigid' control system mainly use aluminum connecting rods. According to theoretical
analysis (Havlín V., 1998), (Potůček R., 2001), composite airplane
control system is about 66% lighter than aluminum one. One of the
objectives of this project is to prove previous statement.
The Czech Republic is one of the biggest producer of 'Ultralight'
(UL) and 'Very Light' (VL) airplanes. Thus, there are available numbers of different plane structures, knowledge, experience, data and
information about the structure design, load etc. Moreover, the weight
in these categories is more than crucial. Therefore, the project will be
focused on this airplane category; however, results of this project such
as e.g. the 'relation stress vs. strength' of the metal end pin connection with composite tube can be used in any other area, even out of
scope of Aerospace Engineering.
Objectives of the Project
Main idea of the presented project is to investigate and improve
mechanical properties of the composite connecting rod with the metal
end pins. The project is focused on the buckling stability of composite connecting rod and strength of the connection of composite tube
with metal end joint. Both project objectives will be investigated practically by performed tests.
The buckling stability of the composite tube depends on the filament, filament composition and used resin. The strength of the connection of the metal end joint with the composite tube depends on the
end joint geometry and type of connection:
● gluing
● winding
● mechanical joining (rivet, bolt)
● combination
Load and Requirements
Increase in airplane speed (dynamic pressure) causes the increase of
the hinge moments of plane control surfaces. However, necessary
deflections of the plane control surfaces are theoretically limited by
the physical potential of the pilot.
Like other parts of airplane structure, the design of the control system has to fulfill requirements defined by the appropriate regulations
(JAR 23, FAR 23, JAR-VLA etc.). Regulations define the minimum
and maximum loads applied by the pilot on the control stick. Aerodynamic load acting on the plane control surfaces (elevator, rudder, aileron etc.) are calculated or measured.
12
Test Load Definition
Chosen airplane category indirectly defines the test load used in the
project. The regulations JAR-VLA defines max. and min. allowed
load applied by pilot on the control stick. However, real load carried
by the control system mechanism depend on the actual 'hinge'
moments acting on the control surfaces. Airplane control system is
commonly designed in such a way, that load transferred by system is
decreasing toward the pilot (Zálešák F., 1966).
One of the first steps of the project solution is statistical research on
the present situation. In this project, collection of data which represent
the load of the UL and VLA airplane control system have been done.
Control systems and design documentation papers of several
today's UL/VLA airplanes have been used for the statistical analysis.
Analysis itself has been focused on the collection of information
about the geometrical parameters of used control rods and their load.
According to regulation, the airplane control system components
are loaded either by the pilot via the control stick, or by aerodynamic
load acting on the control surfaces (Figure 1). The following tables
(Table 1 and 2) represent collected data from some of the current
UL/VLA airplanes. Due to the fact that the data presented are considered as 'intellectual' property of the plane manufacturing companies, the names of some of the airplanes are not published.
Additional data on geometrical properties and load of airplanes
ting rigidity of the whole airplane control system. Due to high control
rods slenderness ratio, the Eulers' buckling respond under the compressive load can be expected. Figure 2 shows four basic instances of
the slender rod buckling.
Buckling of the airplane control system components can be represented by the second instant of rod buckling shown in Figure 2. The
critical load which causes rod collapse can be written as follows
(Kolektiv, 1989):
(1)
(2)
Buckling resistance of the long slender rods is highly affected by the
axial misalignment of the applied compressive load. Allowable misalignment of the control rod end lugs are about 1mm (Mikula J., 2004).
However, any of the misalignment of the end lugs under the applied
compressive load causes the additional parasitic bending load. Thus,
the control rod is subjected to the combination of buckling and bending loads, which significantly decreases rod buckling resistance.
End lug accuracy alignment is one of the problems which cannot
be omitted in the composite control rod design. Alignment of the control rod end lugs relates to the technology of the manufacturing process.
Figure 2: Basic instances
of the slender rod buckling. Source: (Kolektiv,
1989)
Figure 1: Load carried by the airplane control system
Control System Self Frequency
Table 1: Geometrical property and load of control rod loaded by control stick
Besides the rigidity, important parameter of the airplane control system is its self frequency and system response to the aerodynamically
excited oscillation of the airplane control surfaces. Frequency analysis and 'flutter' response of the whole plane structure is required by
regulations. Frequency behavior of the airplane control system is out
of scope of this project. Self frequency of the whole airplane control
system highly depends on the system configuration, bearing, used
material etc. However, in a plane design the frequency analysis of the
airplane control system is crucial and cannot be omitted.
Present Composite Control Rods
Table 2: Geometrical property and load of control rod loaded by aerodynamic load acting on Elevator
control system rods of aileron, rudder, landing gears etc. have been
also collected, however they are not presented here. According to this
analysis, the geometric properties and test load of the control rod specimen are defined. At least three different length of control rod specimen will be tested, max. length will be 3m. Control rod diameter will
be about 28mm and thickness of the rod wall about 1.0 to 1.5mm.
Max. applied static load will be about 3000N.
First successful attempts to utilize the 'composite control rod' in the
airplane structure have already been made. 'Composite control rods'
based on glass filament are in use since 1996. These kind of control
rods (Figure 3) can be found in Czech airplane such as Lambada,
Samba or Seehawk.
The shape of the 'glass' control rod is conical with aluminum end
lugs. Rod itself consists of two halves laminated from unidirectional
glass fabric and consequently glued on the wide welts. In longitudinal
direction is 'glass' control rod stiffened by several carbon rovings.
Control Rod Geometry
Control rods of the plane control systems are usually very slender
long tubular components loaded by compressive load.
Such components lose their axial stability under the relatively 'low'
compressive load. Thus, control rods buckling behavior define resul-
Figure 3: 'Glass' control rod used in planes Lambada, Samba,
Seehawk
13
Presented project is focused on the 'carbon' composite control
rods. First kind of such control rod (Figure 4) is used in the plane
VL-3 (Figure 5).
Figure 4: 'Carbon' control rod used in VL-3 airplanes
Figure 5: The VL-3 airplane
The 'carbon' composite control rods are made by 'winding' technology, so the rod cross section is circular, with constant thickness of
the rod wall. The aluminum end lugs are glued into the composite
tube. These 'carbon' composite control rods are made by Czech company CompoTech s.r.o. This company has great experience with the
carbon winding technology; they are able to manufacture the composite tubular structure even with 0° filament layout in the rod longitudinal direction. This feature is advantageous in the tubular composite
structure loaded by bending, due to high flexural rigidity. In the case
of the airplane control rod, flexural rigidity is crucial due to rod buckling resistance. Thus, presented project is focused on the development
and testing of the composite control rods with 0° filament layout.
Planned tests will be focused on the composite rod buckling resistance, influence of end lug misalignment, strength
of the connection metal end lug & composite tube, influence of the
long control rod accommodation in the sleeve bearings, etc.
End Joint Arrangement
Despite the fact that composite materials offer a number of advantages by their utilization in all kind of structures, the remaining problem
is the distribution of single load acting in one point. In some cases the
problem can be solved in similar way as in the metal structure. However, a number of cases require better 'strength-fatigue' behavior than
composite materials offer. However, the previous statement does not
have to be entirely true, mentioned problem can be considered as the
result of the insufficient manufacturing technology and knowledge of
the 'strength-fatigue' behavior of the composite materials. Connection
of metal end pin together with composite tube represents the application desired by industry; however, the knowledge of 'strength-fatiguedeformation' behavior is missing. The aim of this project is to develop
reliable connection 'composite vs. metal' and investigate the 'stressstrength-fatigue' characteristic of such a connection.
Figure 6 shows several possible types of metal end lug connection
with composite tube. First top left picture represents the glued lug, top
right picture represents the glued and riveted lug, and two bottom pictures represent the winded lugs to the tube.
Presented options of metal end lug connection with composite tube
represent only the preliminary idea; however, not the final one. It is
obvious that during the project solution the design and system of the
connection of the metal end lug and composite tube will develop.
3/2005
End Lug — Composite Tube Joint Analysis
Every joining method presented in the previous section has some
advantages and disadvantages. The glued or adhesively bonded joint
can distribute the required load over a larger area than the pure
mechanical joint, requires no holes, adds very little weight to structures and has superior fatigue resistance. However, the adhesively bonded joint requires the careful surface preparation of the adherend, is
affected by service environments and is difficult to disassemble for
inspection and repair (Lee D.G. et.al., 2004).
The mechanical joint is represented by fastening the components
with bolts or rivets. Since it requires holes for bolts and rivets, stress
concentration, fatigue and galvanic corrosion problems can occur at
the holes. A partial solution of the mechanical joint disadvantages is
its combination with adhesive bonding.
Wound joint like the adhesively bonded joint can distribute the
required load over a larger area and does not require holes. However,
the fourth picture in Figure 6 shows a combination of winding method
with the mechanical joint. The difference and advantage of this type
of end lug connection is that the composite material is during the rod
manufacturing process wound around end lug and lug pins. In other
words, the holes for the lug pins are nor additionally drilled, yet
wound.
Like the adhesively
bonded joint, the wound
joint requires a careful
surface preparation of the
metal end lug; moreover,
the geometry of the lug
'connection' surface is
more complex from the
manufacturing and economical points of view.
Figure 6: Some of the
options of the connection design of composite
tube and metal end pin
References:
[1]
[2]
[3]
[4]
[5]
[6]
[7]
Havlín V.: Konstrukční návrh příčného řízení malého sportovního
letounu; M.Sc. Thesis, Czech Technical University in Prague, Aerospace Department, 1998
Kolektiv: Pružnost a pevnost II; Czech Technical University in Prague, 1989, publication number 6628
Lee D. G., Kim H. S., Kim J. W., Kim J. K.: Design and manufacture of an automotive hybrid aluminum/composite drive shaft; Composite Structure Journal 63, page 87-99, 2004
Marczi T.: Effect of the number of ribs on the aircraft wing gross
weight; M.Sc. Thesis, University of Glasgow, Department of Aerospace Engineering, June 2002
Mikula J.: Konstrukce a projektování letadel, Podvozky, řízení letadla, rozpočty a náklady; Czech Technical University in Prague, 2004
Potůček R.: Příčné řízení malého sportovního letounu; M.Sc. Thesis, Czech Technical University in Prague, Aerospace Department,
2001
Zálešák F.: Konstrukce a projektování letounů — Řízení letounu;
Vaaz Brno, Czech Republic, 1966
14
Tlakové zkoušky pro stanovení materiálových charakteristik
použitých v numerickém modelu FSW
Ing. Petr Bělský / VZLÚ, Plc., Prague
Lately VZLÚ, Plc. — division CLKV 3400 established close cooperation with leading European research and
development centres focused on FSW technology. Within the frame of the activities VZLÚ, Plc. accomplished
a number of material tests for selected Al-alloys. The article gives short information about compression tests carried
out to generating representative input data for the numerical model of Friction Stir Welding.
VZLÚ a.s. — divize CLKV 3400 navázala těsnou spolupráci s předními evropskými pracovišti zaměřenými na výzkum a vývoj
v oblasti třecího svařování (FSW). V rámci těchto aktivit byla uskutečněna celá řada materiálových zkoušek vybraných typů
hliníkových slitin. Příspěvek podává stručnou informaci o uskutečněných tlakových zkouškách. Naměřené materiálové
charakteristiky byly použity pro zpřesnění numerického modelu FSW.
Keywords: friction-stir welding, Al alloys, material properties, tests, tensile, compression,
Hopkinson bar, strain-rate (0.001 s-1 to 3000 s-1), temperatures (22°C to 500°C).
Introduction
Friction stir welding is a relatively new solid-state joining process
patented by TWI in 1991 [1]. It has proven to be a very successful
joining technique for aluminium alloys and has the potential for
welding other common alloys. Superior mechanical properties in
the weld zone are producible in comparison with conventional welding processes. It has been successfully introduced to the aerospace, land and marine transport industries and is now a focus of considerable experimental and modelling investigations.
Because FSW is an emerging technology, there are considerable
risks associated with technology transition into aerospace structures. One strategy to mitigate risks and enable successful process
implementation is to develop modelling and analysis capability for
the new welding method. Numerical models provide for development, optimisation, and qualification of current and future FSW
operations. Fundamental process aspects such as nature of the heat
source, coupling between thermal and mechanical phenomena, and
the development of grain structure and precipitate morphology in
the weld zone are not completely understood. The ability to predict
the effect of welding parameters and tool geometry on the temperature, strain fields, and the resulting weld microstructures is essential for optimising and controlling the FSW process.
Lately VZLÚ, Plc. established close cooperation with leading
European research and development centres focused on FSW technology. Within the frame of the activities VZLÚ, Plc. accomplished
Figure 1 — Hot compression test apparatus
a number of material tests for selected Al-alloys. Materials property data for 3D modelling activity were generated using a combination of tensile, compression and Hopkinson's bar tests. Accurate
numerical FSW model is very data-intensive. Data for wide strainrate range (0.001 s-1 - 3000 s-1) and temperature range (22°C500°C) are needed. Short information about conventional compression and Hopkinson's tests at high temperature levels is given
below.
Hot Compression Test
A servo-hydraulic testing machine Schenck equipped with laboratory furnace ATS 3210 was used to measure mechanical properties
at lower strain rates, from 0.001 s-1 to 1 s-1. All hot compression
tests were performed in accordance with standards ASTM E989/E209-00 and Measurement Good Practice Guide No. 3 from
NPL Materials Centre UK [2].
Very important aspect of the test system is the use of platens,
which are enough to sustain the loads required for hot deformation
and which will not react with the testpiece and lubrication at the
high test temperatures. That is why new platens were made of tungsten carbide (80%WC+20%Co). This material has hardness greater
than 55 HRC and it is possible to use it for tests minimum up to
800°C. Surface of the platens was polished with special diamond
pasta to the brilliant polish. The bearing surfaces of the testing system shall be parallel at all times within 0.02 % unless an alignment
Figure 2 — Stress-strain curves for Al-alloy at strain rate 0.1 s-1,
22°C-500°C (illustrative figure)
15
device of the type described in ASTM E9 is used. In our case new
precise push rods made of heat-resisting steel 17 255 were used.
Complete arrangement of testing system for hot compression test is
shown in Fig.1. Remote sensing of the load train displacement from
an actuator LVDT (linear variable differential transformer) corrected for machine stiffness was used for measurement of longitudinal
strain. External high temperature extensometer Epsilon Model
3548-025M-020-ST attached to the platens was used for control
measurement. Temperature of the platens was monitored using
a thermocouples placed on their bearing surface near to the specimen.
Good lubrication is vital to conducting valid hot axisymmetric
compression tests. It is essential to keep barrelling to a minimum to
ensure that the measured flow stresses are representative of the
material behaviour. Friction effects can be also corrected, but
a value for the friction coefficient must be known. For example an
equation from Dieter (1989) can be used:
For low µ this can be simplified by expanding the exponential term:
Friction coefficients for selected Al-alloys and strain rates were
estimated using the simple ring up-setting test, generally known as
the Cockcroft and Male Ring Test. In the ring test an annular sample of specified dimensions is deformed between platens. From
a measurement of the changes of the internal diameter of the testpiece it is possible to deduce a ”friction factor“, by interpolation of
theoretically calculated calibration curves. If the friction is high the
central hole is reduced in diameter when the ring is subjected to
compression, and increases in diameter when the friction is low.
Because the compression tests were performed at wide temperature range 22-500°C, three different lubricants were used: for 22200°C — teflon spray, 300°C — MoS2 spray Molybkombin
UMF T4 and 400-500°C — BN paste.
The specimens for compression tests were obtained by cutting/machining cylindrical samples from larger sheets. The aspect
ratio of the specimens was 1.5 (do = 8 mm, ho = 12 mm). Machined
surfaces had a surface finish 1.0 µm or better.
3/2005
Stress-strain curves for selected Al-alloy measured at strain rate
0.1 s-1 and temperature levels 22-500°C are shown in Fig.2. Marked changes in material behaviour are evident.
Split Hopkinson Bar Test
The split Hopkinson bar test is the most commonly used method
for determining material properties at high strain rates. The principle of this experimental technique is based on the well-known
one-dimensional theory of wave propagation in elastic bars and
the interaction between a stress pulse and a short cylindrical specimen, which is sandwiched between the input and output bars.
Illustration of the SHPB is shown in Fig.3.
The typical SHPB apparatus consists of a gas gun, a striker bar,
an incident bar (input bar) and a transmitter bar (output bar).
A gas gun accelerates the striker bar (projectile). The resulting
impact with the input bar produces an elastic, compressive pulse
(incident pulse εi). The elastic pulse propagates down the input
bar until it reaches the input bar/specimen interface. This compressive pulse then partially reflects back into the input bar (reflected pulse εr) and partially transmits through the specimen into
the output bar (transmitted pulse εt). These three stress pulses in
the bars are recorded by capacity transducers on the input and
output bars. The amplitude of the incident pulse depends on the
striker velocity. In addition, the length of the incident pulse is
twice the time needed for an elastic stress pulse to travel the
length of the striker bar.
Based on one-dimensional stress wave theory, the values of
stress and strain in the specimen can be determined by using
records of transducers located on the input/output bars. The axial
stress σs(t), strain εs(t), and strain rate εs(t) in the specimen can be
calculated as follows :
Where t = time; A = cross-sectional area of the bar; E =
Figure 3 — Compressive split Hopkinson bar apparatus
16
Young's modulus of the bars; C o = speed of the elastic longitudinal wave in the bar; L s = length of the specimen; A s =
cross-sectional area of the specimen; ε i (t), ε r (t), = incident and
reflected axial strain in the input bar; and ε t (t) = transmitted
axial strain in output bar.
All high strain rate tests were performed in frame of cooperation between VZLÚ Plc. and VUT Brno — Institute of
Manufacturing Technology. SHPB apparatus was equipped
with new bars made of maraging steel VascoMax 350 (Tensile
strength: 2500 MPa, Yield strength: 2400 MPa), new striker
bars 150 - 380 mm and new system for induction heating of
specimen. Because high rates of strain during friction stir welding occur in areas with higher temperature levels (closely
below shoulder), all planned tests were performed at 400°C
and 500°C. It is also maximum for applications of maraging
steel, because above 500°C austenite reversion leads to a dramatic loss of strength and hardness. Special cooling system
protecting bars and capacity transducers was used (see Fig.4).
The input and output bars had lengths 0.8 m. They were
instrumented with capacity transducers for transverse strain
reading. Longitudinal strains ε i (t), ε r (t), ε t (t) in input and output bars were determined by means of Poisson's ratio. Typical
recorded signals from capacity transducers for an Al-alloy at
temperature 500°C are shown in Fig.5. The incident pulse was
generated using a 200 mm striker bar with approximately
16 m/s impact speed. The average strain rate for this test of
1000 s -1 is only ”eyeball estimate“ because strain rate is not
constant in this type of SHPB test. It varied from 900 s -1 to
1400 s -1 . Cylindrical specimen had a diameter and length
9 mm. The orientation of the cylindrical axis of the specimen
was in direction of rolling.
At high strain rates the testpiece temperature can significantly increase and affect results of measurement. That is why
experiments focused on measurement of deformation heating
were performed. Miniature thermocouples made in VZLÚ,
Plc. were used for temperature measurement inside Al-alloy
specimens during Hopkinson's tests at room temperature,
400°C and 500°C.
Temperature rise for tests at strain rate c. 1800 s -1 and room
temperature reached approximately 17°C, which corresponds
to values measured with HgCdTe detector for identical Alalloy at California Institute of Technology in Pasadena [3].
For plastic strains between 0.05 and 0.15 dependence of dissipated heat and flow stress selected Al-alloy during HPBT test
at room temperature was not observed.
Figure 5 — Typical incident, reflected and transmitted pulses from
SHPBT
Figure 4 — Cooling system for HPBT apparatus
The same tests focused on measurement of temperature rise
were performed also for HPBT tests at temperature levels
above 400°C. Material at the temperatures was soft in so far
as no temperature rise was measured. It can be expected that
deformation heating had no influence on results of high strain-rate tests at high temperature levels.
Comparison compression tests at strain rate 0.1 s -1 , 1000
s -1 , performed at temperature 500°C is shown in Fig.6.
References:
[1]
C. T. Dawes, W. M. Thomas: "Friction stir process welds
aluminium alloys: a new friction welding technique allows
easy welding of normally difficult-to-join materials", Welding Journal 75(3): 41, 1996.
[2]
B. Roebuck, J.D. Lord, M. Brooks, M.S. Loveday, C.M.
Sellars, R.W. Evans: Measurement Good Practice Guide
No 3, "Measuring Flow Stress in Hot Axisymmetric Compression Tests", NPL Materials Centre UK, 1997.
[3]
J. Hodowany, G. Ravichandran, A. J. Rosakis and P.
Rosakis: "Partition of Plastic Work into Heat and Stored
Energy in Metals", Experimental Mechanics, Vol. 40, No.
2, p.113-123, June 2000
Figure 6 — Stress-strain curves for Al-alloy at two strain rates at temperature 500°C
17
3/2005
Impeller Design
Návrh oběžného kola odstředivého kompresoru s tandémovým
Doc. Ing. Daniel Hanus, CSc, Ing. Tomáš Čenský, Ing. Ivan Jeřábek /
Department of Automotive & Aerospace Engineering, Faculty of Mechanical
Engineering, Czech Technical University in Prague
In the article an innovative design of the centrifugal compressor rotor tandem cascade is described. The new design
approach is based on the analysis of flow characteristics in the rotating working channels, which results from
previous systematical research on flow phenomena in turbo-machine curved channels carried out at the Center of
Aerospace Engineering. The new design of the impeller flow passage geometry has been made by re-engineering the
existing newly designed centrifugal compressor stage at the Walter Engines Ltd. The rotor tandem blade concept lies
in the separation of the impeller blades into two parts. The first part is formed by the axial-diagonal flow inducer
section of the rotor. This flow part is designed as a highly aerodynamic loaded axial flow compressor cascade. The
second part of the impeller cascade is designed with double number of blades as in the first part of the impeller. Both
cascades are positioned to each other so that the wakes flowing out from the first cascade enter approximately the
center of every second flow passage in the blade to blade channel of the second cascade. The CFD analysis proves
significant improvement in the compressor stage characteristics and shows a potential of possible innovation even of
a present-day advanced centrifugal compressors for the turboprop engines which is now in the stage of development.
Keywords: centrifugal compressor, rotor, impeller blade, cascade, new design.
Nomenclature
b . . . . . .relative coordinate
c . . . . . .absolute velocity
C . . . . .axial flow cascade blade chord
D . . . . .diameter
G . . . . .mass flow rate
KE, KH .exponents of deceleration
l . . . . . .coordinate in the flow direction
n . . . . . .relative coordinate
p . . . . . .pressure
P . . . . . .position ob the place vith maximum camber of the chord
R . . . . .radius
T . . . . . .temperature
s . . . . . .meridional coordinate
S . . . . . .blade spacing
u . . . . . .peripheral velocity
w . . . . .relative velocity
α . . . . . .angle between axial direction and relative velocity direction
β . . . . . .angle between relative velocity and tangential direction
γ . . . . .angle between tangent to the chord of the airfoil and
axial direction
δ . . . . . .angle between relative velocity and meridional plane
ε . . . . . .turning angle of the flow bend in the axial flow inducer
cascade
π . . . . . .Total pressure ratio
η . . . . . .isentropic efficiency of the compressor stage
ν . . . . . .exit flow deviation
ω . . . . .angular velocity of impeller
Subscripts
1 . . . . . .inlet of impeller
2 . . . . . .outlet of impeller
m . . . . .meridional component
u . . . . . .tangential component
Introduction
One of the most important turbojet engine part is a compressor. Its
performance influences not only thermal efficiency of the engine (and
thus operating economy) but also its operating characteristics. Attaining high pressure ratio and maximum acceptable turbine entry temperature (in relation with other working cycle parameters) and paramount efficiencies of engine parts are of the main goals of each development project. It concerns not only design point; the compressor
should be efficient over wide range of operating conditions. Requirement for high efficiencies at high flight altitudes (at high corrected
speed) is introduced with aircraft engines indeed. From simple construction and manufacturing cost points of view it is desirable to reach
required pressure ratio with minimum number of stages. Requirements for a stable operation and ability to perform fast acceleration
call for a compressor with sufficient surge margin over all operating
envelope. The above mentioned problem lays stress on importance of
compressor design methods and their continuous improvement. The
state of art in the domain of centrifugal compressor stage design in the
Czech Republic is one of most advanced worldwide. A very productive indirect design method was formulated, developed, tested and
validated during past two decades in the Aeronautical Research and
Test Institute and later in the Walter Aero Engines Company in Prague by Vaněk and Matoušek (1), (2).
At the Department of Aerospace Engineering and later at the
Department of Automotive and Aerospace Engineering and the Aerospace Research Center at the Czech Technical University in Prague
systematical research concerning the fundamental and applied research focused on the energy transformation mechanisms in the flow
fields within working channels of turbo-machines and jet engines was
carried out. The practical application of the research led to formulation of fundamental conditions and premises needed for obtaining high
efficiency of these energy transformations either in stationary or in
rotating channels. The first idea concerning new design approach to
18
the centrifugal compressor impeller design (4), (5) was further developed and applied on the newly designed advanced centrifugal compressor stage at the Walter Engines Ltd. (3).
Results of the 1 D design then form basic input data file for the following part of the method — design of the impeller blade geometry
and diffuser geometry.
Thermodynamics of the centrifugal compressor
stage
Inducer design
Basic parameters, which describe compressor operation quality, are
the total pressure ratio and the isentropic efficiency defined as the ratio
of ideal compressor isentropic work needed for given pressure rise to
the real input work of the compression for the given pressure rise
from the inlet thermodynamic state of the air. The thermodynamic states of the air within the centrifugal compressor stage are described by
the mean values of pressure and enthalpy or temperature of the air
and its relative velocity. Typical meridional shape of the centrifugal
compressor stage and characteristic flow sections are defined in Figure 1.
Inducer design procedure is based on the optimization of the inlet area
of the inducer for given air mass flow with regard to Mach Numbers
at the leading edge of the blade.
Optimization represents here a selection of such inlet velocity
distribution corresponding to minimum Mach number of the relative
air velocity at the inducer tip.
Exducer design
Specified parameters of the centrifugal flow compressor stage like
a total pressure ratio π (defined as ratio of total pressures at inlet and
outlet), and rotational speed nK are main input data to the iterating
computational procedure leading to the determination of optimum
peripheral velocity u, and blade angle at the outlet.
Diffuser design
Fig. 1: Centrifugal
compressor flow
path — station
numbering
On Figure 2 the T-s diagram of the thermodynamic state curve in the
air flowing through the centrifugal compressor stage within the flow
sections is shown.
3c 4c 5c
T
T5c
2c
5s
4s
3s
2s
c
p1
5is
p5
c
T5is
Fig. 2: The curve of the mean thermodynamic state of the air in the
flow path within the centrifugal
compressor stage in the entropy
The basic diffuser geometric parameters are calculated by a dimensional design method considering the diffuser as a single part without
taking into account that actually it consists of the vaneless part, vaned
part and outlet casing. Diffuser losses are estimated by means of
selected polytropic coefficient.
Calculation of the mean streamline spatial geometry
The indirect design method for the impeller blade geometry definition has been used. This method was developed by Vanek and Matousek [1]. The method gives the basic geometric parameters of the mean
streamline of the impeller flow path and the meridional shape of the
impeller blades.
The method is based on the equation of motion for non-viscid compressible relative steady flow in the rotating system of coordinates
which is described by general vector equation:
(1)
The equation of motion for the system of coordinates: stream-wise
coordinate l, normal hub to shroud coordinate n and bi-normal coordinate b (Fig. 3) can be expressed in analytic terms:
(2)
0c 1c
0s
T1c
1s
s
(3)
Primary centrifugal compressor stage design
Primary centrifugal compressor stage was designed using the inhouse indirect design method. The design consists of three main steps.
First step is based on the one dimensional calculation of the basic
geometrical parameters of the stage that gives a data for obtaining
a complete definition of the impeller blade geometry and diffuser flow
path.
The intention of the one dimensional design is to define mean thermodynamic, kinematical and geometric parameters of the basic stations of the air flow path trough the compressor stage as described in
Figure 1.
Input data file needed to conduct the 1 D design calculation includes the following parameters:
1) Given basic desired performance parameters of the compressor
stage
2) Data obtained by optimization procedures
3) Data based on the experience gained from experimental research
Design procedure consists of three main parts: impeller inducer
design, exducer design, and diffuser design
(4)
Fig. 3: Spatial shape of
the mean streamline and
system of coordinates
19
To solve this system of equations it is necessary to choose the course of the pressure gradients in l, n and b directions as a function of the
streamwise coordinate l.
To minimize the secondary flows due to transversal pressure gradient in the relative flow field in the inter-blade flow channel the
assumption of the zero pressure gradient in direction of n along the
mean stream surface is taken.
If the peripheral and meridional components of the absolute velocity
are defined the equation (4) with respect to the velocity diagram is
then
(5)
The velocity diagram is determined by next figure
Fig. 4: Velocity triangle
To calculate the radius of the curvature Rm of the mean stream in the
meridional plane the velocity diagrams in the flow path must be
known.
As the peripheral velocity u is known at any point of the mean stream surface then the velocity diagram can be determined if the relative velocity vector is known. So the vector of relative velocity must be
prescribed by some efficient formula. The absolute value of relative
velocity is then according to experience given by equation:
(6)
The angle β (measured from peripheral direction):
(7)
The equations (5), (6), (7) define the space shape of the mean streamline. Its shape can be changed only by means of variation of deceleration exponents KH and KE for boundary conditions taken from
preliminary one dimensional design of the centrifugal compressor
stage. When all the thermodynamic parameters are known at any station of the mean streamline it is possible to determine the meridional
shape of the flow path through impeller by using the equation of continuity. Then the mean plane of blades is given by the set of radials
which intersect the mean streamline.
Impeller primary design
Using the above described design method the centrifugal compressor
Fig. 5: Primary design of
the impeller
3/2005
stage was designed for design working point parameters shown in
next table.
Pressure ratio of the compressor
stage
Isentropic efficiency of the
compressor stage
Mass flow rate [kg/s]
Impeller rotational speed [rpm]
Total temperature raise [K]
Preswirl [ ° ]
Inlet axial velocity [m/s]
Mach number of the inlet
relative velocity
Angle of the absolute exit
velocity [ ° ]
Circumferential velocity at the
exducer tip [m/s]
Mach number of the outlet
absolute velocity
Geometric angle of the blades at
exducer tip [ ° ]
Inner diameter of the inducer
[mm]
Outer diameter of the inducer
[mm]
Exducer tip diameter [mm]
Exducer height [mm]
Number of main blades/splitter
blades
Specific speed
4.475
0.824
4.301
37600
186.22
0
200.0
1.261
20.20
526.1
0.9574
45.0
65.0
183.8
267.2
14.45
14/14
127.98
Tandem blade design philosophy
The flow mechanism within the centrifugal impeller inter blade channels is very complex. Different forces act on the air. First of all there
are the mass forces as Coriolis force normal to the relative velocity
vector and centrifugal forces given by rotation of the impeller and by
the curved trajectory of the flow within the channel. Other forces
acting on the flow are contact static pressure and friction pressure forces and inertia force due to the acceleration of the flow. The flow field
within the channel as for given rotational speed of the impeller and the
flow field at the intake affected mainly by the geometry of the impeller inter blade channels. The working ability of the impeller and its
efficiency are conditional on stable boundary layers at all flow rounded surfaces on blades, hub and shroud. As the Coriolis force is proportional to the relative velocity magnitude, the low momentum fluid
flow in the boundary layers is gradually transported by the transversal pressure gradient from the pressure side of the blades to the suction side creating a wake and important non uniformity of the outlet
velocity distribution. Mixing losses then reduce the efficiency of energy transformation within the impeller and affects also unfavorably the
work of the diffuser. The history of the boundary layer evolution from
the inlet section of the impeller plays important role in the quality of
energy transformation. As the channel is relatively long there is possibility to reduce the boundary layer thickness at the blade suction
side by division of the blade into two separated parts, each part of the
blade geometry optimized for maximum stability of the flow and considerably sufficient load.
Two parts of the blade then form tandem cascades placed close
behind each other and positioned in a way that the wakes flowing out
from the first one enter approximately in the middle of channels of the
next cascade. The design of the tandem blade comes out from the overall design of the impeller using the method described above. As the
impeller blade cascade is formed by basic long blades and splitter short
blades, the tandem blade separation plane was chosen in first approximation at the leading edges of the intermediate blades. In this case the
first cascade represents axial flow compressor cascade and the second
cascade represents centrifugal compressor axial/radial cascade.
First axial cascade geometry is designed then as highly loaded cascade with supersonic relative velocity at the tip part of the blades. The
curved camber of the airfoils of the designed blades is of a circular
arc. The airfoil profile shape is then defined approximately as a double arc airfoil so the maximum thickness lies at 50% of the airfoil
20
chord. The thickness of the blade varies from the root to the tip in the
same way as the thickness of the inducer blades. The inlet camber
angle β1 of the axial cascade blades and its variation from the root to
the tip of the blade is identical to the angle of the inducer, so the incidence angles of the axial cascade and inducer are the same. The airfoil turning angle and its variation from the root to the tip of the blade
is defined by the optimum conditions for axial flow compressor cascades empirically identified by Howell.
Fig. 9: Alternative tandem
blade cascade impeller
design
(8)
As for the chosen position of the plane of the axial flow blade cascade separation the length of the chord of the airfoils is given, the equation defines the optimum turning angle ε of the flow in the cascade of
given density S/C.
The exit deviation angle ν of the flow in the axial flow cascade is
described by empirical Constant equation
(9)
The airfoil camber angle is then calculated by the iteration process
using the equation
(10)
Angles α1, α2, γ1, γ2 are measured to the axial direction and have a negative sign. Calculated values of γ2 along the blade from the root to the tip
define the curved cambers of the airfoils from the root to the tip and the
thickness of the airfoils and its variation enables one to define the axial
flow blade geometry.
The outlet flow velocity vectors angles α2 are then calculated using
the equation for every given radius at the out-flowing edge of the blade.
The second centrifugal flow axial-radial cascade has the same meridional shape as the primary cascade but it is necessary to modify the
geometry of the inlet part. The angles of attack of the out flowing air
from the axial cascade on the leading edge of the second row of the
axial-radial blades should be close to zero.
Centrifugal compressor stage CFD analysis
The centrifugal compressor stage geometry was designed using the 3D
CAD Unigraphics software. This type of geometrical data gained from
the design process enables one to connect the design directly with manufacturing process on one side and on the other side with the fluid flow
computational process using CFD methods.
The flow fields within the compressor stage air flow passages have
been calculated using three dimensional (3 D) model of the analyzed
flow created with the aid of Unigraphics. This model is then exported via
format Parasolid into program Gambit. In case the transport to Gambit
is not successful the geometrical model in Gambit has to be modified
with the aid of the volume modeler. Computational mesh is generated at
the modeled volume of the flowing fluid in the next step. Relatively simple and fast mesh from tetrahedrons is used but more elaborate multiblock mapped mesh from hexahedrons shows generally less numeric
error within computations. When generating computational mesh it is
useful to create as many domains as possible with more concentrated
mesh for boundary layer, certainly if it is allowed by complexity of the
model. If such domains cannot be generated in Gambit then wall domains are concentrated in Fluent through adaptation of the computational
mesh. Due to axial symmetry the flow volumes both of the impeller and
diffuser are modeled only as segments. This saves great volume of the
computational mesh and more concentrated and thus more perfect mesh
can be selected and it saves considerably also a computer time for the
task solution.
The calculation of the flow fields within the compressor stage was made
on the SGI Altix computer with 24 processors Itanium2 using the Fluent
solver 6.2.16m and quadrilateral and hybrid grid with 365 000 cells in
one computational segment formed by one blade to blade channel within the impeller and the diffuser.
Solution takes approx. 15 000 iterations and the speed with eight
parallel processes was about 1000 iterations per hour.
Preprocessor Gambit output generates computational meshes both of
the stator and rotor, which are directly utilized by program Fluent. After
computational mesh is loaded into Fluent the boundary conditions both
for impeller and stator are established. With regard to high Mach numbers at the impeller and diffuser a compressible environment has to be
selected and thus pressure or eventually mass flow boundary conditions
Fig. 7: Tandem blade design process
Fig. 10: Quadrilateral
and hybrid grid with
total number of 365000
cells in whole computational segment comprising the impeller
and the diffuser and
computed static pressure distribution at the
impeller blades
Fig. 8: Tandem
blade cascade
design process
are considered. Periodic boundary conditions are applied for segments of
21
3/2005
rotor and stator. Unequal number of impeller and diffuser vanes of the
centrifugal compressor stage brings necessity to use computational meshes boundary conditions Moving Reference Frame and Mixing Planes
between rotor and stator.
Interaction between rotor and stator and flow parameters in the sliding
plane of both computational meshes can be investigated in Fluent by
using Model Sliding mesh but only when numbers of impeller and diffuser vanes are commensurable (the same angle of rotor and stator segments is required). If evaluation of the interaction is important this simplification is under certain assumptions acceptable. With regard to high
air compression at the compressor stage the solver Couple is used. Despite relatively great requirements for the computer operating memory
Coupled Implicit solver proved to be practical. If the analysis is carried
out in the range of compressor stable work with sufficient margin a stationary solver setting is sufficient. Air is modeled as compressible and
viscous fluid. Originally a model of turbulence as single equation one —
Spalart-Allmaras was used but now the use of two equations models kε or k-ω is tested. Predominantly with regard to computational mesh size
a wall function is selected for boundary layers. Highest value of the
dimensionless distance from wall y' should be in the range of 60 to 100.
To obtain stable computation process different corrections of the
mesh and of the turbulence model are necessary according to the values of the relaxation residuals. Criterions used for computation completing are as follows: behavior of all residuals (course should be horizontal), absolute value of residuals (less than 10-3) gradient of the mass flow
change at impeller input (zero).
Fig. 12: Calculated compressor stage isentropic efficiency performance characteristic tandem blade versus primary
First results
Fig. 13: Calculated compressor stage temperature increase performance characteristic tandem blade versus primary
First CFD calculations of the flow fields within the primary centrifugal
compressor stage and the tandem blade centrifugal compressor stage
give first results which allow us to compare both cases and with regard
to the internal aerodynamics and with regard to the performance characteristics. On the next figures a comparison of calculated performance
characteristics of both centrifugal compressor stages is shown.
Conclusion
First design alternative of the innovative design of the centrifugal compressor impeller was conducted with the aim to improve the performance characteristics of the primary designed centrifugal compressor stage.
The impeller tandem blade design was applied on the primary designed
impeller in a way to assure the interchangeability of both impeller
designs in the compressor stage casing structure. First results obtained
by the CFD show in the comparison of both cases that the hypothesis of
the influence of new geometrical and aerodynamic configuration of the
impeller blades really improves the efficiency of energy transformation
within the impeller and all centrifugal compressor stage. The lower obtained pressure ratio and total temperature increase is caused by lower circumferential momentum of the air at the outlet of the impeller caused by
the flow field within impeller. It is evident that the geometry of the
second, radial, part of the impeller cascade has to be optimized in order
to get the same pressure ratio as the primary compressor stage.
0 -q
eta - q
0
1
0,95
NTK611_Walter
0,9
0,85
0,8
0,75
0,7
0,65
0,6
0,55
0,5
0,0006
0,00064
0,00066
0,00068
0,0007
0,00072
0,00074
q
dT - q
dT
0,68
NTK611_CVUT
NTK611_Walter
0,66
0,64
0,62
0,6
0,58
0,56
0,54
0,52
0,5
0,0006
0,00062
0,00064
0,00066
0,00068
0,0007
0,00072
0,00074
q
The performance characteristics of both compressor stages, primary
and tandem blade were calculated using the Fluent solver
First results of CFD calculations validate the original premise of the
design philosophy that the tandem blade impeller design can improve
the compressor stage performance characteristics, namely the efficiency
and extent of the stable segment of the performance characteristics.
The alternative tandem blade impeller design will be optimized using
CAD and CFD and then physically tested on experimental stand of Walter Engines a. s.
Acknowledgements
This research was carried out in the Aerospace Research Center supported by the Czech Ministry of Education, Sports and Youth and as
a part of the project supported by the Ministry of Industry and Trade
of the Czech Republic.
References:
[1]
Vaněk V., Matoušek O.: Report on the Research and
Development of the Centrifugal Compressor Stage,
Research Report (in Czech) VZLU (Aeronautical
Research and Test Institute), Prague, 1986
[2]
Vanek V.: The Method of Centrifugal Compressor
Impeller Design and Manufacture; In: Zesszyty Naukove Politechniky Lodzskej, Nr 349, Cieplne Maszyny
Przeplywowe, z. 86, 1979
[3]
Beranek M.: Complex Design Method of the Centrifugal Compressor Stage; Research Report D2-125,
Walter a. s., Prague, 2003
[4]
Jerie J., Hanus D.: Centrifugal Compressor Impeller
for High Pressure Ratio, Czechoslovak Patent No.
AO 233 938, 1985
[5]
Hanus D.: Centrifugal Compressor Research and
Development - Collaboration with the VZLU (Aeronautical Research and Test Institute); Research Report
V88-01, Department of Aerospace Engineering, CTU
in Prague - (in Czech), 1988
5
NTK611_CVUT
NTK611_Walter
4
0,00062
0,7
0
4,5
NTK611_CVUT
3,5
3
2,5
2
1,5
1
0,0006
0,00062
0,00064
0,00066
0,00068
0,0007
0,00072
q
Fig. 11: Calculated compressor stage pressure ratio performance characteristic tandem blade versus primary
22
Užití spektrální dekompozice při snižování hluku vrtulových
letadel
Ing. Tomáš Salava, DrSc, Ing. Marcela Šloufová / VZLÚ, Plc., Prague
Aircraft noise is always composed of many components, which are generated by different noise sources. Spectral
decomposition is one of the many methods, which can be used for identification, and quantification of significant
particular noise components in the overall complex noise signal. In this paper the principle of spectral decomposition
is explained briefly, and its application is shown in two examples of the noise analysis of two different propeller driven
aircraft.
Hluk letadla sestává vždy z více dílčích hlukových komponent, které vytváří různé zdroje hluku. Spektrální
dekompozice je jedna z mnoha metod, které lze užít pro identifikaci a kvantifikaci jednotlivých hlukových komponent v
komplexním, celkovém hluku letadla.
V tomto příspěvku je nejprve stručně vysvětlen princip této metody a její užití je ukázáno na dvou příkladech analýzy
hluku dvou různých vrtulových letadel.
Keywords: propeller driven airplane, noise, spectral analysis, discrete Fourier transform.
1. Introduction
Considerable resources have been invested in noise reduction of airplanes, mainly since the first, very noisy turbojet, and turbopropeller airliners came into operation. Substantial noise reduction was
reached especially by the second generation of turbofan propulsion
already in Eighties. In the following years it has turned out that
further noise reduction, not only of large airplanes, is more and
more difficult. However, further abatement of the noise emitted by
aircraft of all categories is necessary. This article is aimed primarily on the noise of propeller driven airplanes. The described method
can be used also in other applications.
In civil transport aircraft the turbofan propulsion is most often,
now. Propellers are still used e.g. with airplanes for ultra heavy
transport, and on the contrary, with smaller and especially light and
microlight airplanes. With ultra heavy aircraft the turbo-propeller
propulsion is taken for more advantageous in certain respects, than
turbo fan. Propellers driven by piston engines are the cheapest propulsion for small and light airplanes. In comparison with fans, propellers emit more noise with specific strong, low-frequency components. However, propellers of the advanced design need not be
always the dominant noise source of a propeller driven airplane.
Propeller itself is only one of many noise sources in a propeller
driven airplane. Substantial and sometime dominant noise may
generate the engine, which is also a source of strong vibration.
Vibration generated by the engine is always more or less transferred into the airframe, and transformed into an additional noise,
which is emitted e.g. by the vibrating walls of fuselage. Considerable noise is evidently generated also aerodynamically, mainly on
fuselage, wings or undercarriage. The share of the particular noise
components in the resulting perceived overall noise may strongly
depend on the concept and construction of an airplane.
If the noise of a machine is composed of different components,
it is very useful to identify the main noise sources first, and then to
quantify the contributions of the particular noise sources to the
resulting noise. It has no sense to care for less strong noise sources,
unless the most prominent noise sources are made substantially less
noisy. Therefore the decomposition of a complex noise into its components can be very effective in noise reduction practice. In this
article first some basics of noise perception are recalled, then the
principle of the spectral decomposition method is explained briefly,
and finally, two examples of the spectral decomposition application
in the analysis of the noise of propeller driven aircrafts are shown.
2. Noise and its perception
Our ears are basically the logarithmically scaled receptors [1]. In
this way our auditory system manages to work in an immense range
of sound intensities, from the weakest sounds to e.g. the roar of
a big jet plane. This is also why the logarithmically scaled decibels
were introduced for expressing the perceived ”strength“ of sound.
Regrettably, this property of our auditory system has some not
quite pleasing consequences in noise reduction. Reducing emitted
acoustic power of a noise source to one half causes only 3 dB
decrease in the perceived noise level, which is only a bit above the
threshold of perceptibility.
Accordingly, quieting only one of a two approximately equally
strong noise sources results in a similarly weak decrease in the
perceived noise level. The perceived difference in this case could
be more significant if the two noise sources generate sounds with
markedly different spectral composition. To evoke a subjective
feeling that the loudness of a noise was decreased to one half, it
is necessary to reduce the sound level on about -9 dB, which
means to decrease the acoustic power of the emitted noise nearly
to one tenth. The feeling of how much a noise has been decreased
depends also on the level of the ambient, background noise, and
is more significant, if the most annoying components are depressed.
Our auditory system is also approximately logarithmic in the frequency domain. The perceived difference in the pitch of a two
tones is given not by the simple frequency difference, but by the
ratio of their fundamental frequencies. Therefore logarithmic frequency scales are mostly used in acoustics e.g. in spectrograms of
audio signals. One octave in musical terminology is just equivalent
to the ratio 1:2 on the frequency scale. Our auditory system works
also basically as a spectral analyzer, with spectral resolution approximately 1/3 octave. Therefore the one-third-octave spectral analysis is mostly used in noise control and evaluation practice [2].
Furthermore, the feeling of the loudness of a sound depends also
on frequency. Approximately bellow 1 kHz, sensitivity of our ears
23
3/2005
diminishes considerably toward low frequencies,
Therefore the higher-frequency components are
always considerably more annoying then low,
and the lowest components even need not be
audible at all, if they are not strong enough... In
noise measuring practice this is taken into
account by so called frequency weighting. Most
often used is the ISO standard frequency weighting, denoted by the letter A. Sound pressure
level in dB, which was measured with frequency
weighting A is usually given as dB(A). Beside
weighting A, also weightings B, C and D was
standardized, (D for aircraft noise) but these are
less often used [2], [3].
3. Noise sources identification and
spectral decomposition
Different methods can be used for identification
of particular noise sources in aircraft and other
machines. One of such methods is based e.g. on
measuring sound intensity [4], which also makes
it possible to measure separately sound power,
emitted by different particular noise sources.
Another method is based on using large arrays of
microphones to recognize the directions from
which the particular noise component is emitted.
By special processing of the signals from the
microphones, the particular noise emitting areas
on a machine can be visualized, which can be
especially impressive e.g. on a photography of
the flying aircraft [5].
In comparison with the two mentioned and
similar methods, the usability of spectral decomposition is rather limited. However, it needs no
special measuring instrumentation or procedures, and can be very fast. Only a few short samples of the noise of a tested aircraft are mostly
sufficient to be recorded and processed. The proFig. 1 Basic procedures in the SPAD decomposition program
cessing and evaluation can take several seconds
on any PC or notebook. The method proves especially effective for discrete spectral components in complex spectra. Main inner proextracting the dominant ”tonal“, discrete spectral components. cedures in program SPAD are shown in Fig. 1.
The input data are digitized noise signal, preferably in the MicThese are typical for propeller noise and exhaust noise of piston
rosoft WAV format with 8 or 16 bit quantization and sampling freengines, but also for fans.
To identify reliably which of the recognized equidistant series of quency 22.05 or 11.025 kHz. After the data are read in, it is posdiscrete spectral components belongs to the engine and which to
the propeller only three further data items are needed: number of
blades of the propeller, rotation speed of the engine shaft, and
reduction ratio of the reductor gearbox, if any. With turbo-propeller propulsion usually only the propeller noise is fully periodic and
in the spectrogram the propeller components can be recognized
easily, without any further data.
Extraction of the identified spectral components for further evaluation is basically simple. Specific problems may, however, arise in
more complicated spectra with automatic recognition and correct
assignment of the many particular discrete spectral components.
Not quite without problems is also re-transforming the extracted
spectral components and the spectral ”remainder“, back into the
time domain for so called auralization, which means for listening
and aural evaluation of the decomposed noise components.
4. Spectral decomposition with program SPAD
Program SPAD was written for experimental evaluation of different
algorithms of automatic recognition and assignment of particular
Fig. 2 An example of an input signal time course graph
in an 8 segments format
sible to look over the time course of the entered signal, as it is
shown in Fig. 2. Next, the input data are transformed from the time
domain into the frequency, or spectral domain. For this purpose the
standard FFT (Fast Fourier Transform ) algorithm is used [6].
Data from the input data block are transformed in successive sub
blocks with the length of 8192 signal samples. Up to five partly
overlapping sub blocks can be transformed subsequently and averaged. For further processing cumulated or averaged power spectra
are used. The process can be watched in the next window, the
example of which is shown in Fig. 3. The power spectra of the individually transformed sub blocks are consecutively traced in different colors and finally the averaged power spectrum is added. In
Fig. 3 the final averaged power spectrum is drawn in blue color.
The graph has logarithmic frequency scale, and the power density
is expressed in dB with linear dB scale, as is the common practice
in acoustics. By averaging there can be suppressed accidental, random effects.
Fig. 3 An example of a cumulated and averaged spectra of a turbo
propeller aircraft noise
The aim of the following procedure is to obtain reliable and complete information about the particular extractable spectral components contained in the analyzed noise. In this step first, the salient,
local narrow maxima in the frequency spectrum are detected. Then
the relations in their characteristic frequencies are tested, to recognize harmonic, equidistant spectral series, which could belong to
just one particular cyclic process. Such cyclic process can be rota-
Fig. 4 Auxiliary window of the spectrum preprocessing
24
Fig. 5 Extracted spectral components of the propeller noise [red-L(C),
green L(A)]
Fig. 6 Noise spectrum after the propeller components was removed
[red-L(C), green L(A)]
tion of the propeller or pulsations from the exhaust of the piston
engine, or also e.g. spectral components of the saw-like fan noise.
Different evaluation algorithms have been tested for this purpose. In the last version of the SPAD program, the cepstral technique
is combined with two relatively simple algorithms, which work
directly on the power spectrum. Cepstrum is most simply the Fourier transform of the logarithmed power spectrum [ 6 ]. It can be
well used for finding the periodicity in the spectrum and finding the
equidistant, harmonic spectral components belonging to one cyclic
process. The method works well even if the individual spectral
components are partly masked by other components of the noise.
However it fails if the cyclic process generates tonal noise with
very weak or no harmonics. Best results were obtained by combining different methods.
In Fig. 4 there is shown an auxiliary graphic window with the particular results of the processing and evaluation of the power spectrum same as in Fig. 3. In blue color there is drawn the averaged
power spectrum, in green the same power spectrum with the frequency weighting A. Drawn in red color is the cepstrum, which in this
case has very sharp, extremely high maximum at 5.86 msec., corresponding to the fundamental frequency about 171 Hz.
The light blue vertical lines mark the spectral components,
which were recognized by another algorithm as belonging to the
dominant cyclic noise generating process. They are evidently the
propeller noise spectral components. On the left side there are the
25
Fig. 7 Power spectrum of a small propeller driven aircraft with piston
engine
3/2005
The sample of this noise has been recorded inside the passenger
cabin of a small, two-engine turbo-propeller aircraft, about 30
seconds after its take off, in its fast climbing phase. The aircraft was
equipped by three-blade propeller. As is usual with two engine aircraft, the propellers are on both sides of, and in short distance from
the fuselage. This is the worst possible concepts of a propeller driven aircraft as to the internal noise. In this case, clearly a noticeable reduction of the inner noise would be possible with more advanced propellers, or other less noisy propulsion.
As a second example has been chosen an external noise of a light,
sporting airplane. The airplane was driven by two-blade propeller and
six-cylinder, four-stroke piston engine. The noise sample has been
recorded during the climbing of the airplane from the distance 100 m,
the airplane approx. fifty meters above the ground. In Fig. 7 is shown
again the cumulated and averaged power spectrum (blue) and the
same spectrum with frequency weighting A (green). On the left side
of the window are the nominal frequencies of the found series of the
strongest equidistant discrete noise components.
In Fig. 8 is the same spectrum from which the located dominant
equidistant spectral components were removed. In the upper right
corner of the graph in Fig. 7 there are the computed equivalent
sound pressure levels of the analyzed noise sample with, and without the extracted noise components. These were determined also
as belonging to the propeller. The equivalent sound pressure levels
are 89.3 dB(C) and 80,4 dB(A) for the complete noise, and 77
dB(C) and 72.9 dB(A) without the extracted propeller noise components. In this case also an advanced propeller could make the airplane up to 8 dB(A) less noisy.
5. Conclusions
Fig. 8 Same power spectrum as in Fig. 7 with extracted exhaust noise
components
corresponding cepstral qvefrencies [6] in msec., and the characteristic frequencies recognized by the second procedure. The frequency 174.9 Hz differ somewhat from the fundamental frequency
171 Hz, as it was determined by the cepstral procedure.
In the next step the results from the processing and evaluation of
the power spectrum are used for decomposition of the spectrum or
for the extraction of the selected spectral components. After the
decomposition, it is possible to evaluate the decomposed components of the noise separately, or even to retransform the decomposed spectra back into the time domain. In SPAD, the extracted
spectral components can be retransformed into the time domain by
the inverse Fourier transform (IFT). More advantageous would be
the digital filtering, with the filter coefficients derived from the
spectral data, which is, however, more demanding. [6]
The extracted spectral components of the propeller are shown in
Fig. 5. The noise spectrum remaining after the eight propeller noise
spectral components have been removed is in Fig. 6. Both in Figs.
5 and 6, in the upper right corner of the graphs, there are the respective equivalent sound pressure levels L(C) and L(A), as they
was evaluated with frequency weightings C and A. With the weighting A, the extracted propeller noise components give the equivalent sound pressure level 101.4 dB(A) whereas for the remaining
turbine and aerodynamic noise it is only 87.7 dB(A). The propeller
noise is about 14 dB louder, than the noise of the engines and the
aerodynamically induced noise.
Spectral decomposition is one of the methods, which in analysis of
complex noise signals can be used for identification and quantification of periodic noise components that were emitted by different
particular noise sources. It needs no unique or special instrumentation, is fast, and can be very effective in the analysis and detailed
evaluation of the noise of propeller driven airplanes. In this paper
this has been shown on two examples of the analysis and evaluation of the noise of propeller driven aircraft.
Program SPAD, which is used in the two examples was written
primarily for experimental work, and suits for laboratory semiautomatic processing. Further work would be necessary to ensure high
reliability and accuracy of this method in fully automatic processing. Reliable recognition and correct allocation of different particular spectral components is still not a simple task with very complex spectra. Also high quality auralisation of decomposed noise
components for aural evaluation still needs further work.
References:
[1]
Moore B.: An Introduction to the Psychology of Hearing;
Academic Press, Cambridge 2003
[2]
Barber A.: Handbook of Noise & Vibration Control; Elsevier, 1993
[3]
International standards IEC 60651, IEC 60804 and IEC
61672
[4]
Fahy F. J.: Foundations of Engineering Acoustics (Chapter
5); Elsevier, 2000
[5]
Vaseghi S. V.: Advanced Digital Signal Processing and Noise
Reduction; Wiley Publishers, 2000
[6]
Mitra S. K, Kaiser J. F.: Handbook for Digital Signal Processing; Wiley Publishers, 1993
26
Úvod do problematiky recyklace termosetických kompozitních
materiálů
Miroslav Valeš / VZLÚ, Plc., Prague; Ing. Petr Kachlík / Brno University of
Technology
The article deals with problems of recycling of thermosetic composite materials used in aeronautical or other
industries. This contribution is focused on recycling methods and possibilities to re-use the processed materials.
Článek popisuje problematiku recyklace termosetických kompozitnách materiálů, používaných v leteckém i jiném
průmyslu. Článek je zaměřen na metody recyklace a možnosti znovu-užití získaných materiálů.
Keywords: ecology, composites, plastics, thermo-sets, green design, bio-degradability, waste recycling.
There is no doubt that environmental impacts play more and more significant role in the industry as well as in utility sectors, and that this
trend will continue. These requirements are embedded into the legislative standards increasingly and in their bottom principle predestine
two significant ways of solution:
—
—
GREEN DESIGN, i.e. selection of materials and technologies,
which have no negative environmental impact, or which are biologically removable;
Expansion of waste recycling — where the ”GREEN DESIGN“
is not applicable.
It is evident that the GREEN DESIGN is far from applying to all products within desired scale and this is why the importance of recycling
will growth in the future. However there are certain materials commodities that are more than issue from the viewpoint of recycling. For
example fibrous thermosetic composite materials, in particular/namely materials based on epoxy, polyester, phenolic or other resins, reinforced by glass, carbon, kevlar, etc. These materials are difficult to
recycle and the relevant technologies of recycling itself are not ready
for general industrial use.
However it will be necessary to solve these problems in the near
future, especially in relation with a considerable growth in application
of these materials, which already, due to combination of potential qualities (relatively low density with high parameters of strength, hardness and toughness), frequently replaced conventional metal, wooden
or other materials. Usage forecast for fibrous thermosetic composite
materials in various industry branches also predestines their boom,
especially in the aerospace industry, but also in building trade, railway
and shipping products, sports equipment as well as in many other products of consumer goods. The trend of growing usage of these materials suggests that wastes are going to grow, whether in production
phase of composite part (wastes from manufacturing, chippings, etc.)
or after the end of technical life of products with composite material.
The problems of thermosetic composite wastes have not been
explicitly solved in legislative. These materials are not classified as
hazardous substances in waste catalogue and neither national, nor
European regulations specify their treatment. Nevertheless, certain
general legislative specifications (e.g. Waste treatment plan in ČR including Decree of Czech Government No. 197/2003 Sb.) specify
decrease of quantity of waste disposal and increasing re-using materials for particular products (e.g. car-wreck up to 95 %). We are in anticipation of analogous requirements also for wastes from composite
materials, as currently the national legislation in this area is fully harmonised with EU legislation.
There are several ways of composite waste treatment. The simplest
one is waste disposal, which has a big environmental impact and at the
same time do not solve all basic problem. Waste disposal is not in harmony with waste plans ČR and EU as well. Also waste burning is
unacceptable from the point of view of ecology, sanitary and economy. In addition, this makes the future usage of waste as a materials
source impossible.
Therefore the optimal way is the choice from different sorts of
recycling technologies, which enable re-using of composite materials,
or their usage in other manners. There are several relevant technologies, from which the following belong to the most important:
Mechanical reduction — method consists usually of multi-step shearing, crushing, milling or crazing followed by product separation according to element size and granularity. The typical product is composite gravel mixture (fibrous material, resin, fillers,
etc.), usable namely as a secondary material into various filler,
panel, thermal isolation, etc…
Chemical processing — method for recycling thermosetic composite materials by separating the material into the individual constituents by using solvents. This method is specialized to obtaining
fibres (glass, carbon, etc.) as a resulting product of recycling.
There are some disadvantages: the necessary solvents are hazardous to health; time of the process; difficult use for composites
with different types of fibres, e.g. mixture of carbon and glass
fibres (all fibres will create mixture hard to separate).
Thermal processes — make possible relatively high share in followup using and utilization of composite waste. Especially two
types of processes come into account:
●
Fluidized combustion — recycling process with energy and
material utilisation. Composite scraps are processed in a fluidised bed within a stream of hot air. Products are usually carried
away with streaming air to the cyclone separator, where un-fluidized waste is returned back and processed products are removed. Incurred gases are liquidated (with possible energy profit)
in afterburner. Resulting product of recycling are recovered fibres. Their next usage depends on features, such as tensile
strength, hardness, etc.
●
Pyrolysis processing — other type of thermal processing,
where is combustible components in composite waste are heated in pyrolysis chamber without oxygen. Resulting material
breakdown includes solids as fibres, relics of filler, any cinders
and eventually metal or other reinforces (c. 45 to 90 % of
weight), gases (c. 1 to 12 % of weight) and liquids condensable
products (c.15 to 50 % of weight). The pyrolysis gases contain
27
usually CO, CO2 and hydrocarbon gas. Their calorific value is
relatively small in comparison with calorific value of liquid products — pyrolysis oils, which is about 30 MJ/kg and nears to
calorific value of fuel oils — see table 1. The significant raw
material from pyrolysis process is separate fibres. Their using
again depends on their features, which result mainly from pyrolysis temperature. Acceptable application may be moulding
compound, particle composites, secondary filler, etc.
3/2005
on of following equation
[3]
Formula [3] is subsequently enhanced by additional kinetic parameters and chemical mechanism.
The pyrolysis liberates less gas in comparison to classical combustion. GC/MS chromatography (gas chromatography-mass chromatography) is often used for molecular residues weight determination.
This method gives quantities of single chemical elements (percentage
by weight) released by decomposition. An example of results is in
table 2. It also enables determination of weight ratio of individual chemical component, heating power etc. The chemical structure is identified by FT-IR (infrared spectroscopy - Fourier transformation). DSC
(deferential scanning calorimetry) is used for kinetic characteristics
study. It must be pointed-out that these thermo-kinetic processes are
non-reversible and hence it is impossible to obtain parent elements by
thermal conversion.
Tab.1. Calorific values of some resins, used for thermosetic composites, and some more materials for comparing (approximately)
Particular means of recycling was more or less investigated
over the world. The results of research are but rambling and,
with some exceptions, insufficient for general application in
industry, including aerospace. It can be supposed that practical application will be a combination of mechanical and thermal waste processing. The result of recycling will depend
mainly on properties of recycled products, obtained in optimal
conditions of recycling technology.
Within this context the research on recycling by pyrolysis
appears to be perspective, as this method is applicable both
for waste from fibrous thermosetic composite material, and
other type of wastes. The research will include the pyrolysis
process itself, optimization of technological conditions and
resulting properties of recycling products. Also the analytical
simulation of this process will be of great importance.
Analytical simulation of pyrolysis requires knowledge of
exact chemistry of the material
(epoxy, polyester resin with combinations of glass, carbon or aramide
reinforcements). It is a important to
know mechanism of chemical bonds
decomposition and characteristic
constants for this process, types of
hardeners and hardener agents included in the resin.
Principle of resin decomposition is
shown on the pictures 1. and 2.; including structural formula for one epoxy
precursor — Bisfenol F (further, Bisfenol A and C are used as well ) and for
its thermal decomposition into a less
complex structure.
All kinetic studies utilize the basic rate equation:
Fig.1. Chemical structures of Bisphenol F type epoxy resin cured with
DDM
Fig.2. Decomposition of chemical structures
Source from: Chemical recycling of glass fiber reinforced epoxy resin
cured with anime usage nitric acid, Weirong Dang, www.sciencedirect.com, Polymer 46 (2005) 1905-1912
Tab.2. Example of total yields of gases delived from the pyrolysis of
epoxy resin with glas fiber reinforced in relation to final pyrolysis temperature
[1]
where α is the conversion of degradation (for polymer material is usual to assume that f( α ) = (1- α ) n , where n is the apparent order of reaction), t is the degradation time and k is the
reaction constant given relation
[2]
where A is the preexponential factor (min -1 ), E is the apparent
activation energy (kJ mol -1 ), T is the absolute temperature (K)
and R is the gas constant (8,3136 Jmol -1 ). The basic is soluti-
Reference:
[1]
Cunliffe, A., M.: Recycling of fibre.reinforced polymeric
waste by pyrolysis, www.elsevier.com; Pyrolysis 70 (2003)
315-338
28
Ing. Milan Chvojka, Ing. Josef Fabián / VZLÚ, Plc., Prague
Improving microaccelerometer´s parameters is a never ending innovation process. The article describes the blocks of
the microaccelerometer that have been or are going to be upgraded in the near future, as compared to the model
used on the MIMOSA micro satellite. Separately are described improvements on mechanical and electronic parts of
the microaccelerometer. The changes and their reasons are listed first, then more details are presented. The article
shows the results of analyzed innovative variants and examples of the newly developed blocks.
Zlepšovaní parametrů mikroakcelerometru je nekonečným procesem. Článek popisuje bloky mikroakcelerometru, které
byly inovovány nebo budou inovovány v nejbližší době, vzhledem k modelu, který byl použit na satelitu MIMOSA.
Samostatně jsou popsány inovace mechanické a elektronické části. Nejdříve jsou změny v konstrukci vyjmenovány a
jsou uvedeny důvody, které vedly k potřebě těchto změn. V dalším jsou změny podrobněji popsány. Článek ukazuje
výsledky analýz jednotlivých variant zlepšení a příklady nově vyvinutých bloků.
Keywords: MAC microaccelerometer, improvements, reliability, redundancy, lifetime, stability,
sensitivity, resolution, MIMOSA project, locking mechanism, failure.
1. Main tasks in the microaccelerometer MAC
innovation
Improvement on electronics parts, mechanical technologies and mathematical modeling induces the possibility to enhance the microaccelerometer parameters. Also the requirements of future new projects
lead to new design of some blocks of MAC-03 microaccelerometer.
Generally, every new generation of MAC must be more stable, more
sensitive, must have higher resolution. After a failure of locking
mechanism on the MIMOSA project, it is very important to know the
microaccelerometer reliability, its possible redundancy and life-time.
used materials were bad. Therefore we returned to the lever construction of the locking mechanisms. The latest locking mechanisms
are more compact, those used inside MAC-01 and MAC-02, but their
function principle is the same. Figure 2.1 shows the drawing of new
locking mechanisms placed on the sensor.
Figure 2.1
2. List of main changes and their reasons
2.1 Changes in the mechanical part
We are making three significant changes in the mechanical part of
microaccelerometer MAC. First is a new housing of the device,
second is new mechanical interface and third is the redesign of the
sensor locking mechanism.
The necessity of the new housing follows from the fact that the
mechanical construction of the old model of the microaccelerometer
had low own frequency and insufficient rigidity. Due to this fact the
printed circuit of GENER-HSK board bulged too much during vibrations. Our goal is to reach the first eigenfrequency of the microaccelerometer housing higher than approximately 100 Hz.
The mechanical interface of the old MAC-03 was statically doubtful and too complex. It conveyed the forces from spacecraft into the
microaccelerometer construction. New interface is three-point structure and has a sufficient single failure tolerance. Each of the three
connecting points is equipped with three bolts.
The most complex reconstruction is effected in the locking mechanism block. These mechanical blocks fix the cubic proof mass of the
microaccelerometer sensor during lift-off. Free movement of the
proof mass inside the cavity stops is very small, about +20 m. But the
proof mass is manufactured from the quartz glassthat is very fragile.
The locking mechanism of the proof mass mission failed during the
MIMOSA and one of the three has stayed locked. The proof mass
was fixed in one corner of the sensor cavity and the whole device cannot measure the accelerations. It was only possible to measure the
stability of the fundamental electronics control circuits of the MAC03. This information is also very important for estimation of the MAC
parameters but the scientific goal was not reached. Locking mechanisms inside MAC-01 and MAC-02 were lever based; the ones inside
MAC-03 were based on the rotating principle. Their construction and
New locking mechanisms (red)
and their position on the sensor
2.2 Changes in the electronic part
Some circuits of microaccelerometer´s electronic part were changed too. Main changes are made in the GENER-HSK electronic
block. This block generates a spectrum of digital periodical signals,
which synchronizes all periodical functions in the whole microaccelerometer. This block also deals with housekeeping signals and
this part was modified too. These modifications increase the measurement stability.
The previous model of the microaccelerometer has two measurement ranges. Switching between these ranges was done by switching polarization voltage between 10 and 5 volts. This switching
was done inside the position control block by subminiature bistable
(flip-flop) relays. ESA norm enjoins use of this relays in new space
devices. Therefore it was necessary to reconstruct this control circuits and use new recommended suitable components.
Also the power supply block of microaccelerometer needs redesigning. The old one was composed of commercial DC/DC converters with primary voltage 6 V. This voltage was with advantage
used in the MIMOSA satellite, but it is not usual in other international projects.
Power supply generates most of noise and disturbances inside
the microaccelerometer. Necessary reconstruction gives the
opportunity to develop special multilevel externally synchronized
power supply with low level of output noise and higher efficiency.
29
The last change in microaccelerometer electronics will be
made in CPU control block. Previous MAC had CPU control
block based on SAB80C166 processor from Siemens. This
processor is not space qualified and is uselessly complex and
powerful for such an application. Therefore reconstruction of
MAC CPU control block is underway, based on use of Xilinx
Field Programmable Gate Arrays with space qualification.
3. Description of the main changes and
improvements
3.1 Main changes in the mechanical part
Previous development of mechanical part of the microaccelerometer was accomplished in small private firms purely equipped with software modeling tools. CLKV centre in VZLÚ
has powerful SW tools for mathematical modeling of mechanical systems. An analysis shows that a sufficient high first
eigenfrequency could be reached only by using a honeycomb
structure of the MAC-04 cover. Other changes in mechanical
construction were focused to mass reduction and reconstruction of the locking mechanisms, described in the item 2.1.
Mechanical interface of MAC-04 was simplified with respect to previous model of microaccelerometer. Now it is the
three point interface with screw redundancy. This solution
minimizes the forces acting from spacecraft on the MAC-04
mechanics and on the other hand this solution allows better
thermal connection between MAC-04 and spacecraft. Use of
three screws in each connecting point increases the single failure tolerance of the interface. The properties of interface
change very little if one screw breaks down.
3/2005
during the design Due to the thickness increase during the
design, the dimensions of the new microaccelerometer model
grew. Material of plate is duralumin. The purpose of calculations is to obtain a panel with 1-th eigenfrequency above 100
Hz.
3.1.4 Procedure of optimalization
After an introductory analysis critical locations were determined, which are predisposed for oscillations on lower frequencies. Big attention was paid to the cover. Experienced with
ribbed milled structure — this construction ensure a conductive connection and so creation of Faraday cage and heat transfer without bigger problems — we worked on this type of
construction. At the beginning there were simple FE models in
MSC.Nastran, models of plates with surface elements
CQUAD with four nodes (linear elements). After that more
complicated models were formed in environment Catia.
Thanks to improved communication between Catia and Patran
the transmission is troublefree. To these models there were
applied quadratic three - dimensional elements CTETRA,
which is tetrahedron with 10 nodes. This element was chosen
after comparative analysis, in which results were compared
from surface and three — dimensional models.
The rib design structure with sufficient first eigenfrequency
was too heavy. For this reason we have focused on a honeycomb structure.
Figure 4.3.1
3.1.1 Modal analysis of MAC-04 cover
There are special requirements for constructions that are
designed for space. Space environment has some specialities,
structures are subjected to space radiation, vacuum, electromagnetic field, micrometeorites. In addition, each system passes through several lifetime stages, must be shipped onto the
launching establishment, set at it, lift-off, separation of sections, aerodynamic cover, disconnection from launcher, work in
space. In the case of return on Earth the satellite is designed
for entry to the atmosphere, breaking and stopping. Every
stage has its specific loads requiring high reliability on the
whole system.
There are various requirements and regulations. One task is
to develop a universal construction satisfying all of them. This
part concerns modal properties of microaccelerometer MAC04 cover.
One of the final
proposals, 1. wave shape
3.1.5 Conclusions of this part
It is demanded that the entire mass should not exceed 5 kg.
Another restrictive factor is the first eigenfrequency, which
has to be above 100 Hz. From principle of function, microaccelerometer demands that the conductive connection of all
structure around measuring element should be created as
a Faraday cage. Equipment will not be placed in a pressurized
area, so it will be subjected to influence of vacuum. From that
follow requirements on materials and construction. Further
requirements are related to form and sizes, from which then
follows limits to thickness of cover too.
This part of project meant a significant step in MAC-04
design. It has been shown that future development will take
direction towards honeycomb construction, because engaged
mass and stiffness requirements cannot be met with milled
ribbed construction. About honeycomb constructions it is
generally known that they are lighter and stiffer, nevertheless
it means to solve also special problems concerning transmission heat, electrical conductivity, vacuum, production technology, etc.
We have to point out some simplifications made: , this
model leaves out damping of air, in addition calculation only
deals with 180 x 180 mm plate and not construction on the
whole.
Data transfer from environment Catia to MSC.Patran is now
almost troublefree. For solving Patran is more suitable to
solve problems with FE method but not convenient enough for
very complex models. So in such cases it is better to use Patran-Catia combination.
3.1.3 Parameters setting
3.2 Main changes in the electronic part
To be able to compare the results of computations a plate
180x180 mm was chosen with constant thickness for various
calculation runs The thickness limit was changed several times
Main changes in the electronic part of the MAC-04 were made
in the GENER-HSK block. The GENER-HSK block generates
a spectrum of synchronization digital signals, which synchro-
3.1.2 Requirements on construction
30
nized all periodical processes inside the microaccelerometer.
This synchronization is necessary to minimize the disturbances
between microaccelerometer blocks. The most sensitive block
in the microaccelerometer is the proof mass position detector,
POSDET. This circuit works with signal about 1 A for the
whole input range. It is extremely sensitive to external disturbances. Position detector needs for its proper function digital
clock 230 kHz and synchronized small sine voltage with the
same frequency and stable phase. All synchronization frequencies are derived from this position detector main frequency.
Sine voltage for POSDET must have high spectral cleanness.
New developed synchronous sine voltage generator has 10
times better ratio of high harmonics than the old one. It is also
newly equipped with circuits for amplitude stabilization and
phase control of generated sine wave voltage. All these new
features allow increase of POSDET stability, resulting in increase of whole MAC-04 stability. Figure 3.2.1 shows the
spectral density of the new sine generator output.
chronized and their efficiency is relatively small. Solution of
this problem is complex, but it is necessary to develop special
multilevel power source with external synchronization. Moreover, this power supply must have very limited output noise.
Because the power consumption of MAC-04 analogue circuits
is relatively stable in time and power consumption of digital
circuits change very rapidly, it seems to be useful to separate
these two arms and supply them from two separate power
sources. This solution minimizes the influence of changes in
the digital arm into the analogue circuits.
Development of the new CPU control block of the MAC-04
is also in progress. Previous model of MAC had CPU control
board based on the SAB80C166 processor. This processor is
not space qualified and is not perspective for future projects.
For cooperation in ESA projects it is necessary to use the processor or components which have space qualification. In cooperation with firm ASIX, s.r.o., which in the Czech Republic
supplies the Xilinx Space Qualified Field Programmable Gate
Arrays (FPGA) and has great experience with its implementation, we are preparing an entirely new CPU control block
based on this technology. We suppose that FPGA will emulate PIC16 and all peripheral components, except ROM and
RAM memory. Such solution could be easy SW tailored for
different spacecraft, with the same hardware. From this point
of view this solution is very cost effective and flexible for the
future use.
5. Future possible MAC microaccelerometer
utilization.
Figure 3.2.1 Spectral density of the sine voltage generated by the new
GENER-HSK block
The GENER-HSK block includes also the circuits that convert housekeeping signals (for example from temperature sensors) to voltages. Previous model of microaccelerometer use
temperature sensors with current outputs. These sensors contain a little bit of soft magnetic materials and are not suitable
for magnetic oriented missions. These thermal sensors were
replaced by platinum ones, which have better magnetic properties. Previous sensors have current outputs, platinum ones
change their resistance with temperature. Therefore conversion circuits must be also redesigned.
Small modification will be made also in analogue control
circuits for position control of proof mass inside the sensor
cavity. Subminiature flip-flop relays will be replaced by circuits with FET transistors. This solution decreases the amount
of soft magnetic materials inside MAC-04 and harmonizes
used electronics details with ESA norms.
Fully new development is in the process on the power source block of MAC-04. Previous models of microaccelerometers
had their power blocks composed from commercial DC/DC
converters. This solution had several disadvantages. Microaccelerometer MAC needs minimally seven different supply
voltages with very different power consumption in each voltage arm. Whole power consumption of the MAC is about 3
Watts. Small power DC/DC converters are usually not syn-
Development of microaccelerometer in the Czech Republic
began in the Eighties of the last century. Primarily, the microaccelerometer was developed for small regular shape satellites, intended for the upper atmosphere dynamic studies. In
course of time practice and demands of several experiments
show that usage of such a device is much broader.
Microaccelerometers are necessary in technological material experiments in space, to determine the ”quality“ of zero
gravity and zero deceleration. In connection with GPS receivers microaccelerometers allow us to determine the tracking
of the spacecraft with high accuracy. This accuracy is unavailable with single GPS, comparable with laser interferometry
methods, which are much expensive.
These features and a wide range of utilization make the
microaccelerometers a very perspective product, usable in the
future in many space missions.
References:
[1]
Ing. J. Fabián: Analýza konstrukce pláště mikroakcelerometru MAC-04. Příspěvek na interní konferenci CLKV-VZLÚ,
a.s. Prague, (in Czech)
[2]
D. I. Kozlov: Konstruktirovanie avtomaticheskikh apparatov;
Mashinostroenie Moscow, 1996 (in Russian)
[3]
V. N. Gushchin: Osnovy ustroistva kosmicheskikh apparatov,
Mashinostroenie Moscow, 2003 (in Russian)
[4]
Ondřej Klos: Diplomová práce - Pevnostní analýza malého
kosmického prostředku, Letecký ústav VUT Brno, 2005 (in
Czech)
[5]
MSC.Nastran 2005 — Quick Reference Guide
31
3/2005
Safety Analyses of Aircraft Avionic Systems
(on the aircraft level)
Rozbor bezpečnosti avionických systémů letadla
Ing. Jiří Hlinka, Ph.D. / Institute of Aerospace Engineering, Brno University
of Technology
The paper is focused on the safety assessment of avionic systems during airplane's design and certification process
with practical application to the avionic system of a small General Aviation (GA) airplane. Topics discussed in the
paper include: requirements on regulations of software and electronic systems, a practical example of certification
analysis for such systems, available sources of information from producers. The driving force for solving of the aforementioned problems was the development of a new generation of small aircraft in the Czech aviation industry. Recent
aircraft are equipped with rather complex avionic systems (usually with presentation on multi-functional displays). It
is necessary to keep such systems as safe and reliable as possible. The presented work is part of Task D2
”Dependability of aircraft systems and equipment“ within a project by Aerospace Research Centre (ARC).
Příspěvek je zaměřen na analýzy bezpečnosti avionických systémů v průběhu návrhu a certifikace letounů s praktickou
aplikací na avionické soustavě malého letounu všeobecného letectví (GA). Problémy popsané v příspěvku zahrnují:
požadavky předpisů na software a elektronické systémy, praktický příklad certifikační analýzy pro zmiňovanou
letounovou soustavu, dostupné zdroje dat od výrobců. Hybnou silou pro řešení zmiňovaných problémů byl vývoj nové
generace malých sportovních letounů v českém průmyslu. V současnosti jsou letadla vybavena poměrně složitými
avionickými systémy (obvykle s prezentací dat na multifunkčních displayích). Je důležité udržet vysokou úroveň
bezpečnosti a spolehlivosti těchto soustav. Část prací tvoří rovněž součást úkolu D2 v rámci projektu CLKV (Centra
leteckého a kosmického výzkumu)
Keywords: avionic systems, safety assessment, regulations, airplane design, certification.
Introduction
The paper presents work done in the frame of Aerospace Research
Centre, Task D2 ”Dependability of Aircraft Systems and Equipment“.
Part of the work is focused on application of reliability analyses on
Part23 aircraft.
Nowadays, complexity of avionic systems for GA (General Aviation) airplanes is growing. Purpose of this development is to enable
flights in adverse meteorological conditions and at night as well as to reduce
workload of the crew. New complicated avionic systems are even on small
airplanes. Such systems include for
example utilization of multifunction
displays for flight data presentation
(which was not usual in the past). Glass
cockpit and instruments for IFR
(Instrument Flight Rules) flights make
small airplanes more dependent on the
electric/electronic systems than ever before. In fact, those systems
were originally used only in transport category airplanes (for example
Airbus or Boeing). Failure of such avionic systems has catastrophic
consequences for an airplane and its crew (especially in bad weather
conditions). It is also not possible (because of their complexity and
new design) to prove their safety on the basis of similarity to existing
systems.
The paper is focused on
the analysis of state-of-theart avionic systems that include complex electronics
and software. Such systems
have special requirements.
In addition to redundancy,
also several tests are required to prove tolerance of the
systems against indirect
lightning strikes and high
intensity radiated fields
(HIRF).
Pic. 1 — Modern avionic systems in small general aviation aircraft
VUT 100 Cobra (at Aero Friedrichshafen, 2005)
Presented procedures were practically applied during the development of VUT 100 Cobra and EV-55 aircraft.
Requirements of regulations
Design and certification of an aircraft is made in compliance with airworthiness regulations. CS (Certification Specifications) are used for
design and certification in Europe. These requirements are similar to
US regulations FAR (Federal Aviation Regulations). GA airplanes use
CS-23 requirements [1] or FAR-23 [2] regulation and related advisory circulars. Recommendations related to the required levels of relia-
32
Table 1 — Recommendations of the
Advisory Circular AC 23.1309
*Note: A reduction of Software Development
Assurance Levels (SDAL) applies only for
Navigation, Communication, and Surveillance
Systems if an altitude encoding altimeter
transponder is installed. This option does not
apply to CAT II/III operations.
bility and recommended procedures are covered in AC 23.1309-1C
[3].
Furthermore, several industrial documents are used as a source of
recommendations for design and certification of aircraft. The most
important documents are SAE ARP 4754 [4], SAE ARP 4761 [5],
RTCA DO-178 [6] and RTCA DO-254 [7]. Another commonly used
document is RTCA DO-160 [8].
Table 1 shows basic reliability recommendations for small general
aviation aircraft [3].
Requirements for software
Since software has different mechanisms of failures compared to
hardware, we usually do not speak about probability of the failure.
Instead, we use the term Software Development Assurance Level
(SDAL). The term SDAL is more suitable because software is not
subject to stochastic (random) failures. It is typically possible to repeat software failures (under the same conditions, the system fails
again) — errors are built in software during its development (programming).
SDALs are defined in SAE ARP4754 [4] and RTCA DO-178 [6]
(for the utilization in the transport aircraft category). Documents mentioned above define 5 assurance levels:
An assumption of exponential distribution was used for failure rate
estimates of avionic system components. Part FMEA (Failure Mode
and Effects Analysis) utilized SDALs for components that incorporate software — see picture 2.
Complex failure modes (with simultaneous multiple failures) were
analysed using RBDs (Reliability Block Diagrams).
It is simply possible to create MTBF (Mean Time Between Failures) estimate for the overall avionic system from data in the analysis
(again with an assumption of exponential distribution). It can be used
also for maintenance purposes (i.e. estimates of spare parts, etc.).
Few modifications on avionic system were made, based on the
results of the analysis.
A — highest safety level
B
C
D
E — lowest safety level
RTCA DO-178 defines procedures that are required during the development of the software (for different safety levels). Procedures include programming practices, programming languages, etc.
Indirect effects of lightning strikes and HIRFs
Electronic equipment is vulnerable to electromagnetic effects. Such
effects include HIRFs and effects of lightning strikes. To deal with
this issue, several test procedures for the equipment were developed
in the past, mainly for transport aircraft. Test procedures for airborne
equipment are listed in RTCA DO-160 [8]. Chapters 20 and 22 define test procedures for HIRFs and indirect effects of lightning strikes.
Producer of electronic equipment should provide an aircraft producer
with RTCA DO-160 Environmental Qualification Form - a list of tests
performed. This list includes also testing levels/categories that tested
equipment ”survived“.
Practical Application
Above listed requirements and procedures were applied to the avionic
system of a real small general aviation aircraft presently under certification (example airplane can be seen in Pic. 1). Avionic system has
presentation of primary flight data on 2 LCD displays. The most
important indicators have mechanical back-up (artificial horizon,
speed indicator, altimeter, RPM, MAP). Analyses used in the safety/reliability assessment included:
●
●
●
●
●
FHA (Functional Hazard Assessment)
Failure rate estimates
Part FMEA
Functional FMEA
RBDs
Pic. 2 - Simplified example of part FMEA for MFD (with Software Development Assurance Levels ulitization)
Data Sources
Usually only limited field data are available for new avionic systems
(since they have only very limited operational record). Equipment
producers usually provide aircraft producers with design estimates of
MTBF for particular equipment. Such estimates are commonly based
on MIL-HDBK-217 [9] predictions. No further information (i.e. failure modes and their probabilities) is usually available.
If particular equipment is sold for significant time period, limited data
from warranties may also be available (to verify design estimates).
Only few equipment producers provide aircraft producer with more
detailed data. For example, producer of NAV/COM receiver for
VUT 100 Cobra aircraft was able to provide the aircraft manufacturer
with failure rates estimates for several different failure modes. Furthermore, based on customer repair data base, some of the estimates were
verified.
33
3/2005
ved to be very useful.
Conclusions
Suitability of MIL-HDBK-217 for reliability estimates is widely
discussed in the scientific/engineering community. In the case of
NAV/COM receiver, design estimates were rather pessimistic (according to the producer of the equipment).
Range of design estimates from commonly used sources may be
very wide. Table 2 shows range of estimates from MIL-HDBK-217F,
RAC NPRD-95 and RAC PRISM. It is responsibility of the analyst to
choose the best value for the particular aircraft type.
Relation between MTBF and failure rate is usually simplified to the
following formula:
where is MTBF - Mean Time Between Failures (h)
λ
- failure rate (h-1)
It is necessary to make an assumption of exponential distribution to
use this formula.
The paper discussed
requirements and procedures used in safety/reliability assessment of aircraft avionic (electronic)
systems and availability
of data for analyses. Practical application of procedures/assessment was
made on avionic system
of real small general aviation aircraft. Finally, the
paper provides the reader
with short information
about the year 2005 activities related to task D2 of Aerospace Research Centre project (only a part carried out on Brno University of
Technology).
Dependability has an increasing importance in the aerospace
industry. Reason for that is an increasing number and complexity of
electric/electronic systems built in aerospace products. Reliability
analyses are now often used even in general aviation aircraft category
(this practice wasn't so usual in the past). Application of new procedures had already been justified in the transport aircraft category
where it led to the increase in safety.
Table 2 — The range of failure rate estimates for selected components
(from different sources)
Summary of activities within ARC
Above-mentioned analyses are closely connected to
task D2 within the project of Aerospace Research
Centre. This task is focused on the research in the
area of dependability of aircraft systems, especially
for aircraft certified under CS-23 regulation. Practical safety/reliability analyses made in cooperation with aviation industry are source of
valuable information for the research. Problems solved during the
year 2005 include:
a)
b)
c)
d)
Analysis of avionic electronic system for VUT100 Cobra aircraft
Analysis of pitot-static system for VUT100 Cobra aircraft
Analysis of engine control system for VUT100 Cobra aircraft
Preliminary analysis of electric system for EV55 aircraft
Furthermore, acquisition of reliability software was made in order to
supplement/enhance own capabilities for solution of problems related
to the dependability. Software available on the Institute of Aerospace
Engineering presently includes:
●
●
●
●
●
●
●
●
●
Literature:
[1]
[2]
[3]
[4]
Reliasoft Weibull++ 6
Reliasoft RGA 6 (Software for evaluation of test/field data)
RAC PRISM
RAC NPRD-95C
RAC FMD-97
RELEX Prediction Module + Part libraries
NSWC MechRel (Reliability data, prediction procedures)
Reliasoft BlockSim6.2 FTI
RELEX Markov (Software for RBD, FTA and Markov analysis)
[5]
Above-mentioned software covers most important areas of dependability. Great part of the software is part libraries. Such libraries
offer commercially available data for analyses and already pro-
[9]
[6]
[7]
[8]
CS-23 Certification Specifications for Normal, Utility, Aerobatic and Commuter Cat. Aeroplanes; European Aviation
Safety Agency, Cologne-Germany, 2004, www.easa.eu.int
Title 14 Code of Federal Regulations (14CFR) Part 23 Airplanes: Airworthiness Standards: Normal, Utility, Acrobatic,
and Commuter Category Airplanes; Federal Aviation Administration, Washington, D.C. 7/2002, www.faa.gov
Advisory Circular AC 23.1309-1C — Equipment, Systems,
and Installations in Part 23 Airplanes; Federal Aviation
Administration, Washington, D.C. 3/1999
SAE ARP 4754 Certification Considerations for Highly-integrated or Complex Aircraft Systems; SAE, Warrendale USA,
11/1996
SAE ARP 4761 Guidelines and Methods for Conducting the
Safety Assessment Process on Civil Airborne Systems and
Equipment; SAE, 12/1996
RTCA DO-178B (Software Consideration in Airborne Systems and Equipment Certification), 12/1992
RTCA DO-254 (Design Assurance Guidance for Airborne
Electronic Hardware), 4/2000
RTCA DO-160D (Environmental Conditions and Test Procedures for Airborne Equipment), Radio Technical Commission
for Aeronautics, Washington D.C., 7/1997
MIL-HDBK-217F Reliability Prediction of Electronic Equipment, US Department of Defense, Washington D.C. 20301,
12/1991
34
Aerodynamic Design of V44 Model
Propeller and Aeroacoustic
Characteristics Calculation
Aerodynamický návrh modelové vrtule V44. Výpočet
aeroakustických charakteristik
Ing. Jan Dostál, Ing. Pavel Klínek / VZLÚ, Plc., Prague
The contribution describes the aerodynamic design of a three-blade model propeller aimed at verifying a VZLÚ — V7
new profile series designed for composite propeller blades of light aircraft. The design is derived from V36o successful
composite propeller. The coordinate profiles calculation in ten aerodynamic sections is followed by calculating single
profiles aerodynamic characteristics and partial aerodynamic optimisation. The determination of optimal propeller
blade twist is the first phase and looking for the optimal taper of blades the second. Then the authors go on with
smoothing out geometric measurements and checking calculations of component distribution of induced velocity on
blades radius. The last significant phase of the design consists in checking aerodynamic propeller noise including
detailed description of computational technique and comparison with the noise of the initial propeller.
Příspěvek popisuje postup prací při aerodynamickém návrhu třílisté modelové vrtule, která má ověřit novou profilovou
řadu VZLÚ — V7, navrženou pro kompozitové listy vrtulí lehkých letadel. Návrh vychází z úspěšné kompozitové vrtule
V36o. Po výpočtu souřadnic profilů v deseti aerodynamických řezech následuje výpočet aerodynamických charakteristik
jednotlivých profilů a dílčí aerodynamická optimalizace. Její prvou etapou je stanovení optimálního zkroucení listů
vrtule, druhou etapou hledání optimální štíhlosti listů. Dále následuje vyhlazení geometrických rozměrů a kontrolní
výpočty rozložení složek indukovaných rychlostí po poloměru listů. Poslední významnou fází návrhu je kontrola
aerodynamického hluku vrtule včetně podrobného popisu výpočtové metody a srovnání s hlukem výchozí vrtule.
Keywords: propeller design, aerodynamics, acoustic characteristics.
1. Introduction
In connection with the manufacturing programme of the Czech
Republic aircraft producers that has been changed significantly in the
past 15 years the requirement for new airfoils of light aircraft propellers has emerged. Aircraft propellers produced by VZLÚ were designed for greater speed that is the speed of current light planes and also
for wooden or metal materials of blades. So in 2003 within CLKV-V6
sub-project a profile series of VZLÚ — V 7 (further referred to as
V-7) [1] was designed intended for composite aircraft blades of JARVLA and UL categories. Six basic profiles were designed (fig. 1)
using an optimizing method of modified genetic algorithm developed
in 2002 (CLVK-V6 sub-project again) [2] and supplemented with
a climbing operator in 2003 [1]. But airfoils of composite blades must
be thicker than the profiles of metal blades and the same is true of
wooden blades which
however cannot present
a convenient aerodynamic weakening in front of
the trailing edge. The
optimization has been
done for two regimes —
take off and cruise, both
of them being on positive
sides of profile polars.
Before starting the
model propeller aerodynamic design it was
necessary to define continuous and smooth depenFig. 1 dences of airfoil constants determining the
outlines of single airfoils
[3] on airfoils relative
thickness. For this purpose the interpolation cubic spline-functions
that were interlined with profile constants of six basic profiles have
been used. But simultaneously the profile constants extrapolation was
successfully defined and so relative thickness of the series was enlarged to both sides from the basic range 8 - 20% to the range of 5 - 24%.
For the propeller blade design it was necessary to define transitional
surface between the selected ending airfoil of the series and a blade
hub circle.
2. Model propeller geometry design
Model propeller conception indicates a prototypal propeller designed
for research into its aerodynamic shape or a structural point but not
for lot production.
V44 model propeller was designed in order to verify the quality of
a new profile series. It can be done best by comparing with another
propeller that has older profiles provided its geometry has been preserved or slightly changed only. For this comparison the three blade
composite adjustable and already successfully tested propeller of
V36o was selected. So in the first phase of the design preliminary
geometric dimensions and blades shape of V44 corresponded with the
initial propeller excluding subtle changes of blade tips. The course of
relative thickness of V36o propeller blade from the root to its tip does
not decline monotonously. The relative thickness of the blade end in
decreasing absolute thickness increases slightly. Simultaneously the
ending airfoil camber also increases which is unfavourable from the
point of view of induced drag. On that account the relative thickness
of V44 blade tip is constant. But for purposes of preserving minimal
thickness of trailing edge V44 blade point depth had to be increased
slightly comparing with V36o propeller. The axonometric view of
V44 blade including airfoil sections is in fig. 2.
For the purposes of aerodynamic calculations the blade has been
divided standardly by ten aerodynamic sections. It was necessary to
specify some aerodynamic characteristics for each airfoil section(s).
35
3/2005
Fig. 2. The Blade of the V44 Type
Propeller with VZLÚ-V7 Airfoils
Y
Y
X
Z
X
Z
3. Blade airfoil aerodynamic characteristics
Airfoil aerodynamic characteristics are the dependence of upward
pressure, resistance and aerodynamic moment coefficients on the incidence angle. Their specification is possible only by either measuring
or calculation. For V44 propeller the combination of both methods
has been chosen.
In 2004 six prismatic models of V7 series basic airfoils were produced in VZLÚ workshops. In the same year three of them the relative
thickness of which was 8%,10% and 20% were measured in VZLÚ
laboratory of high velocity at Palmovka under the conditions of
Mach's numbers corresponding to the propeller operating regime. The
results of 2D measuring [4] have shown suitability of designed airfoils
of light planes composite blades and furthermore they make possible
revising calculation results of eight profiles of relative thickness less than
25% with Swshl7 solver in the range of incidence angles of -8° +10°.
Due to interpolation of measured results the incidence angle range was
extended from - 13° to + 19°. Aerodynamic characteristics of two thicker root airfoils were obtained due to interpolation between calculated
airfoils and circular cylinder for that the characteristics are known from
the professional literature [5, 6]. Swshl7 solver was developed in 2002
(CLKV-V6 subproject again) [7].
The possible use of Fluent commercial software and its 6.1.22 version
was also considered for the calculations of airfoil aerodynamic characteristics [8], but finally this idea was dropped for the following reasons:
1) Significant errors of calculated CD drag coefficient contrary to
the experiment,
2) Longer time of the calculation.
The number of regimes that were necessary to count proves how
huge amount of time the calculations demanded. In 1° step 19 regimes fall on one profile airfoil. The total number of regimes for 8 profiles makes 8x19=152 regimes.
The graph in fig. 3 shows CL lift coefficient dependence on the
incidence angle for 10 airfoils of aerodynamic sections made with
a propeller blade from its root to its tip. For the reason of transparence the curves are displaced up. Each curve the serial number of which
is bigger than 1 by 0.4 above the curve whose serial number is less by
one unit. The graph in fig. 4 shows again the dependence of CL lift
coefficient but on the blade radius this time. Mutual curve displacement makes 0.2 here. Balancing CL coefficients on radius is important
from the point of view of induced velocity components calculation.
Characteristics of CD drag coefficients are drawn in figures 5 and 6.
The lines are displaced again, this time by 0.02 in both graphs. The
dependence on incidence angle in fig. 5 shows only little perceptiveness of thick root profiles to the change of the incidence angle, which
corresponds to the large radius of leading edge. CD coefficients significant imbalance towards the axis of zero angle incidence is typical of
V7 series airfoils whose 2 design points corresponding with take off
and cruise regimes are located on the positive incidence angles side.
Fig. 3. Blade V44 - Airfoils V7 - Cl(alfa)
Fig. 4. Blade V44 - Airfoils V7 - Cl(r)
1-35,7%
2-29,2%
3-24,4%
4-20,4%
5-16,9%
6-13,8%
7-11,3%
8-9,5%
9-8,6%
10-8,6%
-13deg
6
-10deg
-8deg
-7deg
4
-6deg
5
-5deg
-4deg
-3deg
-2deg
4
-1deg
3
0deg
Lift Coefficient Cl
Lift Coefficient Cl
+1deg
2
+2deg
3
+3deg
+4deg
+5deg
+6deg
+7deg
2
+8deg
+9deg
1
+10deg
+12deg
1
+15deg
+19deg
0
-13 -11 -9 -7 -5 -3 -1
1
3
5
7
-1
Angle of Attack alfa [°]
9 11 13 15 17 19
0
130 195 260 325 390 455 520 585 650 715 780
-1
Blade Radius r [mm]
36
Fig. 6. Blade V44 - Airfoils V7 - Cd(r)
Fig. 5. Blade V44 - Airfoils V7 - Cd(alfa)
1-35,7%
-3°
2-29,2%
-2°
0,5
3-24,4%
0,5
-1°
4-20,4%
0°
5-16,9%
+1°
6-13,8%
+2°
7-11,3%
+3°
8-9,5%
0,4
0,4
+4°
9-8,6%
+5°
10-8,6%
+6°
Drag coefficient Cd
Drag Coefficient Cd
+7°
0,3
+9°
+10°
+12°
+15°
+19°
0,2
0,2
0,1
0,1
0
-13 -11 -9 -7 -5 -3 -1
+8°
0,3
1
3
5
7
9 11 13 15 17 19
Angle of Attack alfa [°]
Negative lift coefficients possibly used in the reversal are not included
in the work region of V44 propeller blades because their reversal is
not supposed. The graph in fig. 6 showing drag coefficient dependence on the radius under the conditions of the constant incidence angle
starts only from the angle of α = -3°.
The file of aerodynamic airfoils characteristics appears to be one of
the calculation programme input data population of Afrsx2 propeller
aerodynamic solver developed in 2004 (CLKV-V6 sub-project again)
[9]. This solver programmed according to the vortical method has
been used for the calculations of component aerodynamic optimization and induced velocities.
4. Component aerodynamic optimization
was carried out in two phases. The criterion for both phases was
a maximum of propeller propulsive efficiency in one selected flight
regime that was a cruise flight.
0
130
195
260
325
390
455
520
585
650
715
780
Blade Radius r [mm]
The first phase of component aerodynamic optimization was a specification of optimal blade twist for optimized regime so in other
words finding optimal angles of setting propeller blade in ten aerodynamic sections. Calculation recursion of selected regime with successive change of angle setting-up always in one aerodynamic section
coded in input file of blades sizes makes reaching the aim possible
relatively fast because affecting optimalized section with setting-up
remaining aerodynamic sections is low. The enlargement of V44 propeller blade total twist by 9° ( ~ 28%) on the contrary to V36o propeller has resulted in increasing effectiveness by ~0,4%. Larger blade
twist is after profiling the second substantial difference of V44 propeller comparing with V36o initial propeller. The second phase of
V44 blade optimization was seeking optimal blade taper under the
condition of keeping the geometric shape (the equal relative change of
total blade depth). In this phase the absolute and relative airfoil thickness of each aerodynamic section has been changed according to the
empirical relation
Fig. 7. Axial Components of Induced Velocity Propeller V44
24
Take off Vinf=0km/h
Climb Vinf=115km/h
Cruise Vinf=190km/h
Maximal Vinf=237km/h
22
20
18
Velocity Va[m/s]
16
14
12
10
8
6
4
2
0
0
0,1 0,2 0,3 0,4 0,5 0,6 0,7 0,8
Blade Radius r [m]
where ta [m] is the absolute profile airfoil thickness, tr [1] is the
relative airfoil thickness and b [m] is an airfoil chord length (blade
depth). So in diminishing blade taper (chord stretching strain) the
absolute thickness increases and the relative thickness decreases and
vice versa. Due to the influence of airfoil chord changes even Reynold's numbers of single airfoil (circulation) in the considered regime
have been changed. The change of relative thickness and Reynold's numbers results in the aerodynamic characteristics change of these
profiles. As the check calculation of all characteristics in each step of
this phase would be time-consuming, only a few points (2-3) in the
area of work incidence angle of each profile were checked and each
characteristic was displaced by the average value of the difference
between new and original values of the appropriate coefficient. This
phase has not produced more significant effectiveness improvement
but confirmed optimal blade taper of the initial V36o propeller.
The contribution of both optimizing phases depends on the calculation
accuracy of single airfoils aerodynamic characteristics and the whole propeller with the mentioned solvers.
37
Fig. 9. Radial Components of Induced Velocity Propeller V44
24
16
22
14
20
12
18
10
16
8
Velocity Vr [m/s]
Velocity Vt [m/s]
Fig. 8. Circumferential Components of Induced
Velocity - Propeller V44
14
12
Take off Vinf=0km/h
Climb Vinf=115km/h
Cruise Vinf=190km/h
10
8
6
3/2005
Take off Vinf=0km/h
6
Climb Vinf=115km/h
4
Cruise Vinf=190km/h
2
0
-2
0
0,1
0,2
0,3
0,4
0,5
0,6
0,7
0,8
4
-4
2
-6
0
0
0,1
0,2
0,3
0,4
0,5
0,6
0,7
0,8
Blade Radius r [m]
5. Geometric blade sizes smoothing
appears to be a final phase of the aerodynamic design. It was carried
out by successive change of the second derivation of cubic splinefunctions that were interlaced both in 3D area by points of blade (profiles) leading and trailing edges and by 2D dependence of absolute
thickness of airfoils on the blade radius. Propeller blade sizes generation according to set second derivation of spline-functions and the respective circumferential conditions is guaranteed by Ust 4 smoothing
programme of 1995.
6. Courses of induced velocity components along
the blade length
check-calculated by Afrsx2 solver in four flight regimes are drawn in
figures 7, 8 and 9. For purposes of single components comparison
they are drawn in the same scale. Axial components in picture 7 are
dominant; in zero forward speed the axial component reaches about
25 metres per second at the blade tips. In the case of engine cooling
it is very important for axial components sizes not to drop to negative
values in the root parts of blades. As it is obvious in pictures 7 and 8
the graphs of axial and circumferential components have the smoothest courses in the travelling regime. It complies with the course of
angle of setting single profiles optimalized for maximum effectiveness in cruise flight. Induced velocity radial components in picture 9
are not practically influenced either by angles of setting or aerodynamic characteristics of single airfoils and they have a smooth course in
all regimes. The radial component reaches the greatest absolute value
in take off regime again. Radial components negative values give evidence of the current contraction behind the propeller and radial and
the appropriate axial components ratio is equal to current line slope
tangent at the meridial level.
7. Method of calculating propeller noise
characteristics in distant acoustic field
The noise causes a significant strain of human organism. But it increases due to increasing motorization and mechanization of working
activities so it is necessary to reduce single sources intensity. For purposes of reducing propeller aerodynamic noise it is necessary to map
acoustic pressure field and determine its dependence on constructional characteristics of the given propeller. The programme for the calculation of propeller acoustic pressure value in the distant acoustic
-8
Blade Radius r [m]
field has been elaborated according to the calculation proposal by J.
Šulc. The calculation method has been described in report [10].
The calculation algorithm has been processed for Excel spreadsheet. The reason is the widespreading of the mentioned software and its
programming simplicity. The other advantage is an immediate interpretation of results in a form of graphs without any need of other
graphic programmes. The method is based on the calculation of Lm
acoustic pressure levels of the first five harmonic components of the
propeller rotational noise with fm frequency in the acoustic field points.
The field is defined with the angle ϑ of sound generation and r distance. The forward motion influence is being considered in sound generation. The values of acoustic pressure levels are calculated by means
of equations dealing with mth harmonic frequencies of both basic
components of propeller rotational noise that are caused by:
Steady force rotation on propeller blades, the component is marked
as LmS
Blade rotation with the given continuous variable cross section, the
component is marked as LmT
LS and LT acoustic pressure total levels including fm all selected harmonic frequencies follow from LmS and LmT components. Sum of
both LmS and LmT makes the total value of p acoustic pressure level.
Despite the fact that infinitely many possibilities of harmonic components can be included into the calculation, it is quite useless to carry
out calculation for more than first ten harmonic components. Even
first five harmonic components are quite sufficient. So the patterns of
acoustic pressure level calculation embrace the influence of harmonic
components number, air density, propeller blades number, propeller
radius, harmonic components frequency, sound speed, thrust distribution and twisting moment along the blade length. The calculation also
includes Mach’s number influence, sound generation angle and the
distance of an observer from the propeller.
This way of solving propeller acoustic pressure calculation is done
by using some simplified conditions. As an example could serve not
including Doppler’s effect influence and sound intensity loss due to
surroundings impact. The original algorithm [10] did not contain an
acoustic filter of A type [11] used in real measuring. For this reason
the original programme draft has been modified slightly and also this
38
Acoustic pressure levels generated by V36o propeller
Method by [10] without influence of acoustic filter of A type
Fig. 10
Fig. 11
70
70
65
65
60
60
55
55
50
Statistical method (LASt), method by [10] (L), influence of acoustic filter of A type (LA)
50
LS
LT
LASt
L [dB]
L [dB]
L
L
LA
45
45
40
40
35
35
30
30
25
25
25
45
65
85
105
Úhel vyzaøování
125
145
165
25
45
sound filter effect on the resulting value of acoustic pressure level has
been included. The purpose of this modification is to reach the best
result approximating the real measuring.
For calculation purposes the following values of a basic parameter
are necessary to be known: propeller blades number, boss diameter,
propeller diameter, propeller speed in the given regime, TAS flight
speed, air density in the given flight level, sound speed corresponding
with the flight level, harmonic components number (five is enough),
the observer distance from the propeller centre, sound generation angles, values of coefficients of thrust and twisting moment in dependence on the section position on the propeller blade, surfaces of single
sections.
The current programme works with ten propeller sections totally.
The number of section can be increased, possibly the user could be
allowed to have a choice of the sections number. The same is good as
to the number of harmonic components. On the basis of results it is
possible to determine single harmonic components proportion of propeller loudness and propeller geometric shape impact on generating
noise of aerodynamic origin. The programme utility was verified on
the basis of the report results only [10]. The results have not been
confronted with the real measuring but just with the statistical method
results of propeller propulsion noise determination [12].
8. Comparison of V44 and V36o propellers noise
Values of acoustic pressure levels for V36o and V44 propellers were
calculated with the aid of the described programme. The results were
confronted with the statistical method results. [12] Both propellers are
trifoliate having the equal diameter. The calculation of the acoustic
pressure level has been carried out for the identical values of speed,
flight speed and flight height. The number of harmonic components
was five and also was identical. The difference between propellers lies
in thrust distribution and twisting moment along the blade length. Surfaces of single sections are also different.
The calculation of acoustic pressure value characteristics in
65
85
105
125
145
165
Úhel vyzaøování  [o]
[o]
a distant propeller noise field was restricted to determination of L
acoustic pressure level dependence on ϑ generation direction in 40° <
ϑ < 150° interval in r = 300 m constant reference distance.
Generation angle interval was determined with respect to the calculation accuracy. The bigger interval is no use because the obtained
results would be distorted and would not correspond with measuring.
As it is obvious from resulting graphs the noise reaches its maximum
at a level of propeller disk rotation and just tightly behind this level.
Further the effect of both propellers geometric shape is also obvious.
According to the calculations V44 propeller proved to be less noisy
than V36o propeller. However, this effect becomes evident solely in
case of the calculation method according to [10]. The statistical method is not able to incorporate it. According to the statistical method
both propellers generate the equal level of sound pressure. Within the
scope of direction angles of 85 - 150 degrees it is possible to speak
about the equality between the results of statistical method and the
method according to 10.
9. Conclusion
The aerodynamic design of V44 propeller appears to be the first part
of this propeller complex design including composite blades construction, strength, stiffness and dynamic checking by way of finite
elements method. The construction of the hub and mechanism of propeller blades adjusting on the ground will be the same as in case of
V36o initial propeller. A new airfoil series with airfoils thinner in their
trailing part slightly reduces V44 propeller noise in a distant field
comparing with V36o propeller but it can reduce stiffness and worsen
dynamic blades characteristics as well. In this case it would be necessary to rerun the described aerodynamic design but under the conditions of changed relative thickness along the blade length.
The production and flight condition tests of V44 model propeller
are planned for 2006. The acoustic level calculation of the aircraft
noise according to L 16/1 direction, appendix No 7 will be added to
the method of propeller aeroacoustic characteristics calculation. This
39
Acoustic pressure levels generated by V44 propeller
Method by [10] without influence of acoustic filter of A type
Fig. 12
70
65,000
65
60,000
60
55,000
55
50,000
50
LS
LT
LASt
L [dB]
L [dB]
L
L
LA
45,000
45
40,000
40
35,000
35
30,000
30
25,000
25
25
45
65
85
105
125
145
165
25
45
65
85
105
125
145
165
addition will make it possible to calculate the noise of a plane during
take-off, flight over a measuring point and approach. The purpose is
to get closer to the real measurement results.
Fig. 14
Fig. 13
70,000
3/2005
Acoustic pressure levels generated by V36o and V44 propellers
References:
[1]
Dostál, J.: Návrh profilové řady pro kompozitové listy vrtulí
lehkých letounů; Zpráva CLKV-VZLÚ {CLKV-VZLÚ
Report}R-3540/03
[2]
Dostál, J.: Optimalizace návrhu vrtulových profilů modifikovaným genetickým algoritmem; {in} Letecký zpravodaj
3/2002, VZLÚ, Praha, listopad 2002, s. 54-58
[3]
Dostál, J.: Tajný analytický model obrysu profilu; Zpráva
divize leteckých vrtulí VZLÚ R-LV-2002-24
[4]
Dostál, J.: Vrtulové profily V7-8%, -10%, -20%. Vyhodnocení měření; Zpráva CLKV-VZLÚ {CLKV-VZLÚ
Report}R-3657/03
[5]
Jaňour, Z., Podzimek, J., Hacura, V.: Základy aerodynamiky
a mechaniky letu; Nakladatelství Naše vojsko, Praha, 1953,
s. 98
[6]
Hošek, J.: Aerodynamika vysokých rychlostí; Nakladatelství
Naše vojsko, Praha, 1949, s. 388-391
[7]
Dostál, J.: Solvery a kontrolní výpočty obtékání 2D aerodynamických profilů; Zpráva CLKV-VZLÚ {CLKV-VZLÚ
Report} R-3452/02
[8]
Klínek, P.: Srovnávací výpočty obtékání vrtulového profilu
V4-10% solverem Fluent; Výzkumná zpráva CLKV-VZLÚ
{CLKV-VZLÚ Research report} V-1834/05
[9]
Dostál, J.: Výpočtové programy aerodynamiky vrtulí podle
vírové metody radiální nosné úsečky; Zpráva CLKV-VZLÚ
{CLKV-VZLÚ Report} R-3656/04
70
65
60
55
50
LASt
L [dB]
V36o L
V36o LA
V44 L
V44 LA
45
40
35
[10] Šulc, J.: Výpočet charakteristických hodnot ve zvukovém poli
leteckých vrtulí V 510 a V 530; Zpráva Ústavu termomechaniky ČSAV T-396/89,1988
30
[11] Madejewski, B.: Aeroakustika - Základy teorie a aplikace na
konstrukci letadel; VUT v Brně, Brno, 1986, s. 76
25
25
45
65
85
105
125
145
165
[12] Aerospace Information Report, AIR 1407, Květen 1977
40
Zatížení vodorovných ocasních ploch při manévru
Ing. Ivo Jebáček / Institute of Aerospace Engineering, Brno University of
Technology
The article deals with tail load measurements as taken on the KP-2U Sova airplane during manoeuvres. The tail plane
load values were evaluated from strain-gauge data.
Příspěvek se zabývá měřením složek zatížení vodorovných ocasních ploch při manévru na letounu KP-2U Sova s
využitím tenzometrů. Po instalaci tenzometrů a kalibrační proceduře byla do letounu instalována potřebná aparatura a
provedeny měřicí lety. Z naměřených dat bylo vyhodnoceno zatížení vodorovných ocasních ploch při manévru.
Keywords: KP-2U Sova airplane, flight tests, horizontal tail surfaces, load measurements.
1. Introduction
The measurement of loads on aircraft in flight is required for a variety of purposes such as in research investigations, structural integrity
demonstration and development flight testing. If we think about the
possibilities of flight loading measurements as global characteristics
in a given structural cross section, there are two basic principles:
Fig. 2. Installed horizontal tail
1) Pressure distribution measurements
2) Strains distribution measurements
The main disadvantage of pressure distribution measurements is the
necessity of having knowledge of how pressure is spread along the
structute, which requires considerable amount of measurement points
and some changes in the structure itself. Furthermore, another loading
calculation from measured pressure distribution is required. In case
the strain gauges are used, it is possible to achieve good results with
small amount of measurement points, and the measurement equipment can be set so that individual loading components can be scanned
directly during flight.
Fig. 1 KP-2U Owl aircraft prepared to measurement
hods. In total, 16 measurements were taken, and matrix of responses
was 4x16. More detailed description was published in.[1].
3. Measurement and comparison of measured data
and calculation
The KP-2U Owl ultralight aircraft was equipped with ESAM Traveller data acquisition and other needed sensors for displacement of the
elevator, load factor, flight speed, flying attitude, temperature, angular speed, pitch and roll. All data were recorded with a sampling frequency of 100Hz. Weighing and fixing of centre of gravity were done
before each flight.
Flights themselves were exwecuted in such a way that after the
aircraft stabilization, the pilot did sudden displacement of elevator
upwards and downwards or downwards and upwards and then back
to the point of departure (see JAR-23.423 Manoeuvring loads). This
manoeuvre was repeated with different starting speeds. Fig.3. shows
ta typical measurement record.
In our case we were interested in stabilizer's loads of the KP2U Owl ultralight aircraft (see fig.1). This stabilizer was intended to
the laboratory purpose and for flight testing too.
2. Placing strain gauges and calibration
Strain gauges were installed on the horizontal tail, so there was the
security of the transmission of bending and torque moments and
movement forces. This was achieved with 22 attached strain gauges
that were connected into half and full bridges to achieve maximum
sensitivity to the given load.
Then the structure was calibrated. The calibrated coefficients and
equations for the calculation of individual load components, including
presumable mistakes of measurement were set with least square met-
Fig.3. Typical record of a few parameters
After individual flights, responses from installed strain gauges were
recalculated into final loading by manoeuvre with the assistance of
calibrated coefficients.
The calculation was done on the basis of input data from measurement such as speed, multiple of acceleration, centre of gravity, weight
41
Fig.4. Manoeuvre - speed 160 km/h with maximal positive elevator displacement 2.0° and negative one -5.3°
and displacement of elevator. Further theoretical aerodynamic characteristics were accepted just as they were used at the load calculation of the horizontal tail during aircraft development.
The following figures show the comparison of load between the
horizontal tail and theoretical calculation for starting manoeuvre
speed 160km/h (see Fig.4) and for starting manoeuvre speed
120km/h (see Fig.5.). Loads increments are shown as absolute ones
without balancing load.
Fig.5. Manoeuvre - speed 120 km/h with maximal positive elevator displacement 4.2° and negative one -6.9
5. References:
[1]
Jebáček I.: Measurement of Horizontal Tail Load by Strain
Gauges; in Letecký zpravodaj, Praha, 2004, ISSN: 1211877X.
[2]
Jebáček I.: Calibration of strain-gage installation in stabilizer structures for the measurement of flight loads; Grant
project FP 390038, TU Brno, 1999.
[3]
Skopinski T.H., Aiken William S.: Calibration of straingage installation in aircraft structures for the measurement
of flight loads; NACA Report 1178.
4. Conclusion
Measurements of many parameters during flight and subsequent calculation of desired quantities were carried out successfully. It is necessary to be aware of unmeasured values that enter into the calculation
(efficiency of horizontal tail inertia moment of aircraft, etc.), which
affect final computed load to a considerable extent.
3/2005
Použití neuronových sítí při analýze plynové cesty
turbinového motoru
Ing. Jaromír Lamka, CSc / VZLÚ, Plc., Prague
Today many methods are available for gas turbine flow path analysis. Some of them are very simple but yet very
useful, since they give an indication of the compressor capacity with almost no calculation effort. The state of the art
today is the heat and mass balance models, which are more sophisticated. Recently an Artificial Neural Networks
(ANN) have been applied to gas turbine engine diagnostics. In the future, the ANN-based flow path analysis system will
probably, to some extent, replace the heat and mass balance model-based systems, or become a complementary tool
for monitoring and performance analysis of gas turbine engines.
V současnosti je používáno mnoho metod k analýze plynové cesty motoru. Sofistikované modely vycházejí ze stanovení
termodynamické rovnováhy v jednotlivých sekcích motoru s ohledem na mechanické ztráty a ztráty radiační. K tomu je
však nutné znát různé konstanty, které jsou odlišné i pro motory stejného typu. Velice totiž záleží na výrobních tolerancích
dílů, z nichž je motor sestaven, a na utěsnění plynové cesty konkrétního motoru. Z toho vyplývá nutnost individuálního
přístupu ke každému motoru. Právě vzhledem k faktu, že jednotlivé motory stejného typu se svými parametry vzájemně
liší, pak se použití neuronových sítí jeví jako velice vhodné pro monitorování stavu motoru a pro jeho výkonovou analýzu.
V případě neuronových sítí není třeba určovat žádné koeficienty, které se dokonce mohou v celém provozním režimu
měnit. Stačí pouze u konkrétního motoru na zkušebně určit, jaké kombinaci vstupů (učící matice) odpovídá konkrétní
kombinace výstupů (cílová matice). Zmíněnými maticemi je pak navržená neuronová síť "učena" do okamžiku až je
odchylka mezi výstupní maticí sítě a cílovou maticí v požadovaném limitu. V tomto okamžiku je neuronová síť připravena
k rutinnímu (provoznímu) použití. Velice snadno lze pak učit, jak postupně probíhá degradace parametrů hlavních sekcí
motoru a kdy je, nebo bude, nutný zásah údržby.
Keywords: gas turbine engine, diagnostics, neural networks.
Introduction
Engine condition monitoring is an effective complex way to improve
safety as well as reduce operation and maintenance costs of gas tur-
bines. To keep track of the health of various components that make up
a modem aircraft engine, a large number of monitoring and diagnostic techniques have to be applied. Among them, the gas path analysis
Nomenclature
AANN .Auto Associative Neural Network
ANN . . .Artificial Neural Network
b . . . . . .bias vector of the second layer
c . . . . . .bias vector of the first layer
FF . . . .Fuel Flow rating
GPA . . .Gas Path Analysis
hr . . . . .relative humidity
Mk . . . .torque
m . . . . .number of outputs
nG . . . . .rotational speed of gas generator shaft
nP . . . . .rotational speed of power turbine shaft
n . . . . . .number of inputs
pbar . . . .barometric pressure
p2 . . . . .outlet pressure of compressor
Tbar . . . .barometric temperature
T2 . . . . .outlet temperature of compressor
T4 . . . . .inter turbine temperature
W1 . . . .weighting matrices of the first layer
W2 . . . .weighting matrices of the second layer
X . . . . .input vector of the Network
Y . . . . .output vector of the Network
is a kind of fault diagnostic techniques that can be used to isolate and
quantify gas path faulted components of gas turbine engines. Some
features of the gas path analysis (GPA) are the capability to identify
the component responsible for the loss of performance, detect multiple faults and quantify the deterioration affecting individual components.
Performance diagnosis of major gas path components using the
GPA can be carried out by independent parameters, such as component efficiencies, mass flow parameters, etc.
Recently artificial intelligence and especially the artificial neural
network techniques have been applied to gas turbine engine diagnostics. The neural networks have inherent features that make them particularly suited to diagnostic tasks [1].
42
The Multi-Layer Feed Forward Networks (see Fig. 1) have been
used widely for gas turbine diagnostics because of their ability to be
trained by supervised training techniques.
The supervised training methods require inputs and desired outputs
called training data to adjust network weights. The training data are
fed to the ANN and the network provides some outputs. The network
errors could be defined by the differences between network outputs
and target values. The errors can be used to adjust the estimated node
interconnection weights. The steps are repeated until the error lies
below an acceptable level.
The following points are useful in the design of a multi-layer feed
forward network for gas turbine diagnostics applications:
a) Input Layer: The input layer can receive data through engine
instruments during training and real-time application. Via the input
layer the inputs pass to next layer of network to finally produce some
outputs. The number of nodes for the input layer is strongly dependent on application. For gas turbine diagnostics, the number of nodes
should be the same as the number of measured parameters.
b) Output Layer: The output layer produces the estimated output
by performing a variety of defined mathematical function through the
network. For gas turbine diagnostics the ANNs have been developed
following two applications: sensor failure detection and component
failure detection. Sensor failure detection usually employs networks
which are Auto Associative Neural Networks (AANN) (see Fig. 2).
AANN has two distinguishable characteristics from multi-layer networks; AANN has a bottleneck layer and always the number of output and input layer nodes are the same. For fault detection and isolation purposes, one node can be used in output, while the value of output can detect failures or isolate faulty components.
Fig. 2. A typical
ANN structure
Artificial Neural Networks
The presented method is dealing with equations representing some
kind of physical model for the phenomena to be modelled and studied. An alternative approach is to use the ANN, for mapping the system behaviour using the statistical relationship between system inputs
and outputs. Since sensor failure and degradation is a source of increased uncertainty for all different methods, an ANN-based sensor validation method could be used to increase the reliability of all measurement-based system analysis methods.
Artificial Neural Networks are relatively crude electronic models
based on the neural structure of the brain. The first steps to develop
ANN came in 1943 when Warren McCulloch and Walter Pitts [2]
published a paper about how neurons might work. They modelled
a simple neural network with electrical circuits to validate their theory. Later researches stopped because of limitations in learning processes. To overcome these limitations, in 1982, John Hopfield developed
a useful device which was not based completely on a brain model and
it helped to develop lots of fast and practical networks and today, neural network applications can be found in many places.
Fig. 1. Feed Forward
network structure
c) Hidden Layer: The hidden layer enables the network to extract
more information from input data also it makes the neural network
tool a non-linear technique. The number of hidden layers required
depends on the complexity of the relationship between the inputs and
the outputs. Larger numbers of hidden layers will increase the number of parameters and the time needed to train the network. The number of hidden layer nodes should be selected carefully, otherwise
a situation will happen which is called over-fitting. In this situation the
network works well on the training sample but not on test samples.
d) Activation Function: The activation function is usually nonlinear and differentiable for supervised algorithms. The Tan-Sigmoid
or Log-Sigmoid function is typical functions which are used for supervised trainings. The Linear activation function is usually used in output layer of the neural network if absolute value of the output is required higher than 1.
More details about multi-layer feed forward network characteristics are discussed by Anderson and Mc Neill [3].
If the ANN structure is appropriately selected and after sufficient
training, the ANN is expected to represent a generalized model of the
system concerned in the form of simple mathematical relationships.
These models will be able to reproduce and predict the output vector
belonging to the training and to the validating sets of data, respectively, with sufficient accuracy. Their generalization capability, which in
some cases extends to valid non-linear extrapolation, is the focus of
43
their utilization for mapping multi-dimensional input/output data sets
as:
f: Rn →Rm; (y1, y2, ... ym) = f(x1, x2, ... xn)
(1)
where f is a continuous function generated after sufficient training, and
Rn and Rm are input and output data spaces of dimensions n and m.
X=(x1, x2,..., xn) and Y=(y1, y2,..., ym) are input and output vectors,
respectively. It must be noted that the accuracy of the ANN models is
only as good as the accuracy of the data they are trained for.
Function f in (1), which represents the functional relationship in
a feed-forward ANN structure, can be given as:
(2)
where Wl and W2 are called weighting matrices, which are fixed after
training, h is the number of processing elements in the hidden layer,
and band c are weightings of the bias term (+1 in this case); these are
also adjusted during the training procedure.
It must be emphasized that (2) is the generic/static representation of
different systems. The only difference between the ANN mathematical models that are illustrating systems of different types of nature is
in their W1, W2 and b and c values.
3/2005
engine performance model for the same operating point, defined
through an input vector. The set of deltas thus derived is fed to the
ANN for examining whether sensor faults are present or not.
To ensure that Z0 is a reliable estimation of the nominal value,
a model adapted to the particular engine examined must be used.
Adaptation ensures that the effect of engine to engine differences or
the current engine condition are taken into account and are not thus
a source of spurious information. When such a model is used, in the
case of an engine with no component or sensor fault, ∆Z will be equal
to zero. Component or sensor faults cause additional deviations on the
measurements.
Example patterns are shown in Fig. 3. Under the convention used
for this network, pattern (a) is OK, pattern (b) represents faulty engine, healthy sensors, pattern (c) represents healthy engine, faulty sensor, while pattern (d) represents faulty engine and sensors.
Neural Network training and validation
A significant part of using a neural network is training. Training is
achieved through an iterative process where the training data is repeatedly fed into the network and it incrementally improves interconnection weights and biases to match the network data to desired targets.
However, matching network outputs with desired targets do not say
that the network has trained well. Therefore, testing a trained network
with a validation test is essential while a validation test does not include training data. The MATLAB Neural Network toolbox is very
good instrument for this research to investigate the application of neural networks for diagnostics.
Definition of the diagnostic problem
Monitoring of the turbine engine condition is based on a set of measured values on the engine. Although the proposed procedure is of
general validity, the set of measured values must be obtained on the
particular engine type. Only this way the diagnostic of the engine can
be carried out correctly.
Using measured data collected at various operational and environmental conditions, they can generate performance maps using the
ANN model. This model could be used to map the engine performance as a function of the operational conditions.
The performance maps exploiting different system or component
parameters can be generated using an ANN model of simple or complex systems. A comparison between expected and current results
could identify the subsystem degradation and component failure.
The type chosen is a two spool turboprop aircraft engine with a free
power turbine. Fuel flow has been chosen as the cycle parameter setting the operating point, a choice made to give worst case scenarios,
since this quantity is linked to significant measurement uncertainty.
The following functional relationship for generation of performance
maps of turbine engine is:
(T2, p2, T4, nG, nP, Mk) = f(pbar, Tbar, hr, FF)
Fig. 3. Patterns of deltas for different situations: (a) healthy engine
and sensors, (b) faulty engine, healthy sensors, (c) healthy engine,
faulty sensor, (d) faulty engine and sensors
Summary and conclusions
The Artificial Neural Networks are able to detect gas turbine failure,
isolate faulty components and quantify fault index satisfactory accuracy and swiftness compared to GPA techniques. The diagnosis of critical components such as compressors and turbines is difficult without
the monitoring of the gas path. The main benefits of this diagnostics
approach are maintenance time saving and capital costs reduction.
The validation tests indicate that the ANN has high reliability to
detect and isolate failures for degradations more than 0.5% also it has
acceptable accuracy compared to GPA techniques to quantify fault
index.
References:
[1]
Depold, H.R. and Gass, F.D.: The application of expert
systems and neural networks to gas turbine prognostics
and diagnostics; Journal of Engineering for Gas Turbines
and Power, Vol. 121, pp 607-12, 1999.
[2]
Warren McCulloch and Walter Pitts, 1943: A Logical Calculus of Ideas Immanent in Nervous Activity; Bulletin of
Mathematical Biophysics, 5, pp. 115-133.
[3]
Anderson, D., Mc Neill, G., 1992: Artificial Neural Network Technology; A DACS State-of-the-Art Report,
F30602-89-C-OO82, NY.
(3)
In the stage of database construction, a performance simulation program calculates base engine performance and measurable parameters
for various faults. Engine performance degradations are saved at
a file through normalization using the following equation
(4)
where Z0 is the nominal value at the specific operating point (the
value we would obtain from a 'healthy' engine). It is generated by an
44
Měření distribuce tlaku profilu na pozemní pojízdné
laboratoři
Ing. Jan Friedl, Ing. Martin Kouřil, Ph.D., Ing. Róbert Šošovička, Ph.D. /
Institute of Aerospace Engineering, Brno University of Technology
The presented paper engages with description of alternative process to measure airfoil aerodynamic characteristics.
A development of a ground mobile laboratory based on a car is presented. Unlike wind tunnel testing this method is
less money consuming and furthermore tests in conditions of real atmosphere can be conducted implicitly. On the
contrary this situation is complicated by several aspects which influence the measurement. These aspects must be
studied more in future.
At IAE Institute a ground mobile laboratory was built and first test are currently under way. As object for investigation
a "cablo-model" of airfoil MS-0313 was chosen for measurement of surface pressure distribution. Obtained results were
subsequently compared with wind tunnel data of identical model measured in ARTI Prague. Partial results of this
project are shown and discussed in the contribution.
Keywords: alternative aerodynamic experimental setup, car based mobile laboratory, MS-0313
airfoil, aerodynamic characteristics, pressure distribution.
Introduction
Nowadays measurements of airfoils aerodynamic characteristics are
taken mainly by wind tunnel testing. The tunnel itself is always bulky
and rather expensive facility and anyway the final characteristics are
investigated at real peace of work, aircraft prototype. This is the reason
why the flight tests are considered to be more provable, not only in the
field of aerodynamic characteristics. Then there is a logical sense of the
effort in direct testing real aerodynamic characteristics of airfoils.
Real aerodynamic characteristics are mostly measured on the aircraft's flying prototype. Consequently, the flight tests are considered to
be more conclusive, and not only for the aerodynamic characteristics.
The above mentioned reasons support the basic research on the aerodynamic characteristics, especially the design of a new wing section
directly, by flight tests.
Nowadays the flying airfoil testbed at the Institute of Aerospace
Engineering is designed to measure aerodynamic characteristics of
airfoils and for the flight performance and characteristic research in
the area of usage of new technologies, instruments and methods. The
use of flying airfoil testbed makes it possible to include implicitly the
influence of the real flight conditions that are daily encountered in
real operation. Nevertheless, the usage of flying testbeds brings some
problems. Here also corrections on influence of the used aircraft and
ambient air conditions are necessary. The flying airfoil testbed flight
regime must be monitored to determine exactly the conditions of the
measurement. This enlarges the requirements on the precision of piloting and the measuring apparatus on board. Restrictions also include
aircraft´s operational limits and meteorological conditions. With the
aidof this flying airfoil testbed, it will be possible to measure the aerodynamic characteristics of airfoils, leading edge contamination, ice
imitator, manufacturing error, etc...
Model description
The model consists of a rectangular wing section and two circularshape end plates. All components have sandwich construction of
glass-epoxy composite and polystyrene foam. A few ash wood reinforcements in critical positions are used to transfer the local load.
Pressure taps are placed chordwise at certain section. Tubes from
pressure taps are connected with miniature electronic pressure scan-
Fig 1. Structural static test
ners, which are located near left section of the wing. Basic dimensions
of our model are identical to ARTI model for comparison with the
wind tunnel measurement. The rectangular wing section has a span of
1,200 mm and a chord of 600 mm. The circular end plates are 22 mm
thick with 1,080 mm in diameter.
Structural static tests
The basic static tests were performed mainly to obtain an opinion of
the deformation and to check structural integrity of the model. The
computed load was applied on upper wing surface by certain amount
of sand filled bags.
The maximum deformation from bending was 7mm in the middle
of wing span. The maximum deformation angle from twisting was
0.255 degree between two wing end sections.
These values show that the model can be considered as suitable for
our purpose.
Wind tunnel testing in ARTI
The model´s aerodynamic characteristics were measeured in subsonic wind tunnel (the measured area three meters in diameter) at Aeronautical Research and Test Institute in Prague. Pressure and force
measurement of model was done in the wind tunnel up to airspeed
V = 180 km.h-1 including above critical angle of attack.
45
3/2005
Pitot-static probe with the angle of attack and the angle of sideslip
acquisition was mounted onto left end plate.
Measuring process
Fig 2. Wing section model in wind tunnel ARTI in Prague
These measured data were used for comparison with our experiments. Exactly the same model was placed in the ARTI wind tunnel
and on a ground mobile laboratory. In this case, possible differences
in measurement caused by model production can be neglected.
Ground mobile laboratory T-613
The ground mobile laboratory was based on a famous Czech car Tatra T613, which confirms our requests for speed performance and robustness
of construction. Furthermore a high total weight will have positive effect
to stabilize ride with the model and the attachment. The main modifications were realized in a luggage section of the car where a special steel
frame was installed. This frame is used to support the attachment for the
model. The attachment consists of duralluminium tube system, which
will be similar to aircraft mounting (L-13 SE Vivat). The special steel sub
frame compensates the side load. This is located on the right side of the
car and increases width of car approximately twice, so careful driving is
needed. An electromechanical linear actuator is used to avoid twisting of
the model and to vary the angle of attack.
Ground tests were performed at the Brno-Tuřany Airport in an undisturbed early morning atmosphere. The data acquisition itself was performed during steady drive according to inboard speedometer. Each
run gives roughly 40-50 sec. of measured data. Exactly the same
angle of attack was tested in both directions of runway to eliminate
influence of sideslip and to correct the pressure acquisition by pitotstatic probe. An actual angle of attack (influenced by lift of model and
velocity), temperature of atmosphere, and air speed are also collected.
DACU S8256 pressure data acquisition systems together with miniature pressure scanners ESP-32HD (Pressure Systems, Inc.) were
used to collect values from airfoil pressure taps. Each miniature scanner has 32 pressure sensors. Electric power was supplied via onboard
automotive batteries.
Results discussion
So far, two models have been tested. In both cases an airfoil MS-0313
was chosen. The main differences were in number of pressure taps (32
and 64) and slightly bigger diameter of end plates. The second model
has exactly the same dimensions as ARTI model used in their wind
tunnel for comparison purposes. The location of section for pressure
taps was also identical to ARTI model (section location of first model
was the middle of span).
As we expected, the results comparison between measurements in
wind tunnel and measurements on the mobile ground laboratory cannot be considered identical. This is mainly caused by correction absence for proximity of the car and ground. Further research will be primarily oriented for their determination.
Results show significant difference in pressure coefficient distribution
especially on the bottom part. This is caused by 3D flow field affected with
nearness of car. The same effect is shown on lift curves. A change of lift
Fig. 3: Ground mobile laboratory
46
curve slope in comparison with the tunnel data can be described as a local
change of angle of attack also due to flow influenced by car.
Nevertheless, each test of mobile laboratory confirms good repeatability of measurement, so it will be possible to determine suitable correction.
Following figures show measured and computed data for comparison. The figure of lift curves shows also results with surface affected by
rain.
Fig.4 Comparison of pressure coefficient distribution
cp = f(x),tatra α = 6.122 st,Re = 1.22e6
-2.5
tatra top
tatra bot
vzlu α = 5 st,Re = 1.3e6
xfoilα = 5 st,Re = 1.25e6
tatra top,kablo 1,α = 5.867 st,Re = 1.43e6
tatra bot,kablo 1,α = 5.867 st,Re = 1.43e6
-2
-1.5
cp [1]
-1
References:
[1]
[2]
Šošovička, R., Friedl, J.: Aerodynamická měření kablomodelu na pozemním stendu, Zpráva k Výzkumnému
záměru MSM 262100005, Letecký ústav, VUT v Brně,
Brno 2004
-0.5
Šmíd, M., Hanzal, V.: Měření modelu profilu s koncovými
deskami, Zpráva V-1803/04, VZLU a.s. v Praze, 2004
0.5
0
1
0
0.1
Fig. 6 Pressure coefficient distribution, α = 16.75°, measurement runs
22, 23
0.2
0.3
0.4
0.5
x [1]
0.6
0.7
0.8
0.9
1
Fig. 5 Comparison of lift curves
cL = f(α )
cp = f(x)
-5
1.8
top
bot
1.6
-4
1.4
-3
1.2
1
cL [1]
cp [1]
-2
-1
0.8
0.6
Raining during measurement
0.4
0
0.2
tatra Re = 1.22e6
tatra Re = 1.435e6
vzlu cL z cp Re = 1.3e6
1
0
vzlu Re = 1.3e6
2
0
0.1
0.2
0.3
0.4
0.5
x [1]
0.6
0.7
0.8
0.9
1
-0.2
-10
-5
0
5
10
15
20
25
α [st]
Analytical Methods for the Calculation
of Buckling in Composite Sandwich Panels
Výpočet vzpěrné pevnosti kompozitových sendvičových desek
Ing. Martin Baumruk / Czech Technical University of Prague
Keywords: composites, sandwich, buckling, analytical calculations.
Introduction
Composite materials are thanks to their significant advantages increasingly used in aeronautical applications and
other branches of industry. Composite is called a material combined from two and more components, with final properties better than their separate parts. One component serves as a matrix and the other as fibber (reinforcement). In
structural polymer composites, the fibbers (Glass, Carbon, Aramid, Boron) are stiffer and stronger than the matrix
(Epoxy, Bismaleimid, Polyamide). By changing the orientation it is possible to optimize strength, stiffness, fatigue, heat
and moisture resistance, etc. So it is possible to tailor the material to meet specific design needs and get lightweight
material with much higher strength/weight ratio than isotropic materials.
A special example of structure composite material is a sandwich. The sandwich is a layered composite formed by bonding two thin facings to a
thick core (foam, honeycomb etc.). The facings resist nearly all of the applied in-plane loads and bending moments. The thin spaced facings provide nearly all of the bending rigidity to the construction. The core spaces the facings and transmits shear between them and provides most of
the shear rigidity. By proper choice of materials for facings and core, constructions with high ratios of stiffness to weight can be achieved.
47
3/2005
Since composite materials do not exhibit large yielding deformations
prior to failure as metals do and do not allow as much error or approximation in their design it is necessary to examine and predict their
behaviour carefully.
Compressive loadings are inevitable in most aerospace structure, and
therefore it is necessary to study buckling of aircarft's parts such as
panels etc. One classical way of preventing buckling is using stiffeners to reinforce the skin. Sandwich panels offer an alternative to this
since it is possible to achieve a similar bending stiffness and buckling
load without the manufacturing complexity and added joints.
The sandwich structure has many various modes of failure under
compression which can be grouped into two categories: 1 — instabilities such as overall buckling, shear crimping, dimpling of facings,
and wrinkling of facings away from the core or into the core (caused
by insufficient face bending stiffness and core elastic properties) and
2- fracture (face sheets under compression, core under transverse
shear etc.).
There is presented in this paper a study of analytical calculation methods of buckling composite sandwich panels and comparison the
results with FEM analysis.
Figure 1: Basic unit of unidirectional composite material
ABD relations
The stress-strain relation in fibber coordinate system:
(1)
Where Qij is reduced stiffness:
(2)
Transforming Equation (1) to the global x-y-z coordinate system by using the transformation relations and with respect to Kirhoff hypothesis
(when the plate deforms, normal simply translates and rotates and remains straight and normal to the laminate) the stress-strain relations become:
(3)
Where ε° is deformation and κ° curvature of reference surface.
When the stress distribution is integrated through the thickness, a force
per unit length results. This length is in y direction is in case force resultant Nx in x direction.
Figure 3: Definition of force and moment resultant Nx, Ny, Nxy and Mx,
My, Mxy
The force and moment resultants are defined as:
(4)
Figure 2: Possible modes of failure of sandwich composite under
edgewise loads
48
Substituting the stress-strain relations into the definitions of the force and moment resultants, we get a relation between the force and moment
resultants and the reference surface strains and curvatures results. This is the ABD relation — the six by six matrix (so called Hook's law for
laminated composites) and involves a set of parameters that represents the extensional A and bending B stiffness of the laminate, together with
extensional-bending coupling stiffness, which are unique to composite laminate.
(5)
For a single isotropic layer of thickness H and modulus E and Poisson’s ratio ABD is:
(6)
Where
(7)
When the laminate is symmetric Bij are zero and when the laminate is balanced A16 and A26 are zero then ABD relation become:
(8)
Buckling of laminates plate and sandwiches
Buckling of laminated plates simply supported along edges x=0 and y=0
Figure 4: Simply supported plate under in-plane compression before and after buckling
Figure 5 (right): Buckling of plates with increasing initial imperfection
Unlike column buckling, in which a lateral deformation develops along the column length, plate buckling involves two dimensional out-ofplane sine wave deformations.
As the applied load increases, the plate shortens in the load direction and remains flat until a critical load is reached at which the system becomes unstable.
Figure 5 shows that after buckling, the plate is able to carry further load, but at reduced stiffness, which is contrary to the case of column buckling in which the buckling load represents the collapse load.
The knee in load-deflection curve of Figure 5 is only present when the plate is perfectly flat prior to buckling. The knee disappears as initial
imperfection in the plate increase, which also causes a reduction in the load-deflection relationship.
Analysis of plate buckling under in-plane loading involves the solution of an eigenvalue problem in which pre-buckling displacements are igno-
49
3/2005
red. The in-plane loads enter the mathematical formulation of the eigenvalue problem as coefficients of the curvatures rather than as loads on
the right-hand side of the equilibrium equation:
∆=0
K∆
(9)
For a non-trivial solution, ∆ ≠ 0, therefore the solution is found by eigenvalue leading to K=0.
Buckling of specially orthotropic laminated plates
A specially orthotropic laminate has multiple specially orthotropic layer that are symmetrically arranged about the laminate middle surface (to
form a symmetric cross-ply laminate). Neither shear extension (A16, A26), bend-twist (D16, D26) nor bending-extension (Bij) coupling exist.
(10)
The buckling loads are determined from the following differential equation:
(11)
Subject to the simply supported edge boundary conditions
(12)
The boundary conditions are satisfied by:
(13)
Where m and n are the number of buckle half waves in the x and y axes. The solution to the governing differential equation is:
(14)
Since the smallest value forNx occurs when n=1, the critical buckling load is
(15)
Comparison of hand analytical and data sheet calculations and FEM analysis of buckling
We calculated the buckling load of a laminated sandwich panel subject to uni-axial compression. The width of the loaded edge of the panel is
300mm and length is 300mm. The laminate is made up of 6 plies of uni-direction Carbon/Epoxy Fiberdux 913 tape of 0.125mm thickness and
the core is foam Divivicell60. The stacking sequence is [0, 45, -45, foam, -45, 45, 0], The material properties of plies are E1= 138Gpa, E2=
7Gpa, G12= 5,4Gpa, µ= 0,28, t=0,125mm, foam Divinicell60, E= 60Gpa, G= 22Gpa, t= 6mm.
D matrix (from ABD relation) was calculated as:
(16)
The critical load is obtained from:
(17)
Nx = 775 N/mm
To estimate laminate buckling, data sheets are available particulary the ESDU 80023 data sheet for calculation of laminate buckling under conditions of uni-axial in-plane loading. From data sheet
50
(18)
Where width b is measured in metres
Hence
FEM analysis of buckling was carried out by MSC. Nastran/Patran (solution type buckling 105). The critical load for first mode was calculated:
Nx = 775 N/mm
Figure 7: First mode of buckling
Influence of through-the-thickness shear stiffness on buckling
The validity of calculations above is given by assumption that the through-the-thickness
shear stiffness is so large that shear deformation does not significantly affect the buckling
loads.
(19)
If any of the above ratios approaches 0.1 then the through-the-thickness
shear stiffness is such that the buckling loads may be significantly
below the values calculated by previous methods.
The transverse (the through-the-thickness) elastic properties of reinforced composite laminated plates are dependent upon properties of the
matrix material used to bond the fibres together. For many matrix materials, these properties can be very low compared with those in the direction of the reinforcing fibres, with the result that significant deflection
can be caused by through-the-thickness shear stresses which arise in
plate bending.
(20)
References:
[1]
[2]
[3]
[4]
[5]
[6]
[7]
[8]
For the sandwich from example above transverse shear flexibility matrix is:
H 11 = 122.2 * 10 3
H 12 = 0.2378 * 10 3
[9]
H 22 = 116 * 10
[10]
3
And
[11]
So the assumption of large shear stiffness is satisfied.
Figure 6: Buckling curves from ESDU data sheet
L. T. Tenek and J. Argyris: Finite Element Analysis for Composite
Structures; Kluwer Academic Publishers, 1998
N. J. Pagano: Exact solutions for rectangular bidirectional composites and sandwich plates; J. Compos. Mater, 1970
N. D. Phan and J. N. Reddy: Analysis of laminated composite plates using a higher-order shear deformation theory; Int. J. Numer.
Meth. Eng., 1985
A. K Noor: Stability of multilayred composite plates; Fibre Sci.
Technol., 1985
M. Ahmer Wadee and A. Blackmore: Delamination from localized
instabilities in compression sandwich panels; a Department of Civil
& Environmental Engineering, Imperial College of Science, Technology & Medicine, London SW7 2BU, UK, 2000
Structural Sandwich Composites; MIL-HDBK-23, U.S. Department of Defense, Washington, DC, 1968
Rao, K.M.: Buckling Analysis of Anisotropic Sandwich Plates
Faced with Fiber-Reinforced Plastics; AIAA Journal, Vol. 23, 1985
Benson, A. S. and Mayers, J.: General Instability and Face Wrinkling of Sandwich Plates: Unified Theory and Applications; International Journal of Solids and Structures, Vol. 20, 1984
Pierre Minguet, John Dugundji: Buckling and Failure of Sandwich
Plates with Graphite-Epoxy Faces and Various Cores; Massachusetts Institute of Technology, Cambridge
Uemura, M. and Byon: Secondary Buckling of Flat Plate Under
Uniaxial Compressiom Part 2: Analysis of
Stoll, F.: Analysis of the Snap Phenomenon in Buckled Plates;
International Journal of Nonlinear Mechanics, Vol. 29
Anil L. Salunkhe and Prasanna M. Mujumdar: Identification Approach to Estimate Buckling Load of Damaged Composite Plates;
Indian Institute of Technology, Bombay
Aa Mujumdar, P. M., and Suryanarayan, S.: Nondestructive Techniques for Prediction of Buckling Loads - A Review; Journal of the
Aeronautical Society of India
Conclusion
The author presented methods for calculating overall buckling of laminated composite sandwich panels. By comparing the results it can be shown there
is good agreement between analytical solutions and outputs from ESDU data sheet and FEM analysis.
Presented methods of calculation can be used for the first approach in design of sandwich panel under compressive load, but simultaneously it is necessary to keep in mind the limitations of these methods. Buckling load is affected not only by parameters such as axial and bending stiffness (and then by
geometry, thickness of layers, and core, material properties, orthotropic etc.) and boundary conditions but also there is very important influence of imperfections. Imperfections can be both geometrical (flatness) and material. These imperfections have significant effect on buckling load and post-buckling
behaviour. Imperfections can occur during manufacture and service and due to aging and environmental effects; composite laminates are susceptible to
flaws and damage such as voids, porosity, delamination, matrix cracking, fiber breakage, etc. In the case of composite panels designed for compressive loads such flaws and damage may cause degradation of lateral stiffness of the panel and thus of its buckling strength. It is therefore recommended
to compare the analytical solutions with suitable experiments.
51
3/2005
Reliability and Safety Assurance Points of View
Certifikace složitého vybavení z pohledu zajištění spolehlivosti
a bezpečnosti
Ing. Milan Merkl, CSc; Miroslava Nová / VZLÚ, Plc., Prague
Aircraft instrument panel's equipment changes from relatively independent compact instruments towards complex
systems. The remarkable growth of regulations complexity and, in consequence of it, growth of processes and tests
assuring aircraft instrument panel's equipment airworthiness concerns also aircraft certified in compliance with CS-23
specification. The article presents a document system, which is relevant for CS-23 category aircraft and their
equipment airworthiness assurance performed from the viewpoint of safety assurance in the course of the whole life
cycle. The hardware development assurance levels are compared with software criticality levels for given failure
conditions classification. The hardware life cycle phases and software life cycle particularities are described. The
methods and means of airworthiness processes assurance are described through system engineering procedures. A
concept of software tool for CS-23 aircraft certification process and documenting support is recommended.
U vybavení paluby letadla nastává posun z oblasti relativně samostatných kompaktních přístrojů do oblasti systémů.
S tím dochází i pro letadla certifikovaná dle specifikace CS-23 ke značnému nárůstu složitosti předpisů a tím i postupů a zkoušek, jimiž se potvrzuje způsobilost vybavení pro použití na palubě letadla. Příspěvek uvádí systém dokumentů, které jsou relevantní pro osvědčování způsobilosti letadel dle CS-23 a jejich vybavení z hlediska zajištění bezpečnosti požadované v průběhu celého životního cyklu letadla. Pro zavedenou klasifikaci podmínek poruch srovnává
definice úrovní zajištění hardwaru a úrovní kritičnosti softwaru. Popisuje fáze životního cyklu hardwaru a specifika
životního cyklu softwaru. Z pohledu systémového inženýrství shrnuje metody a prostředky zajištění certifikačních
postupů a doporučuje princip softwarového nástroje využitelného pro podporu průběhu a dokumentace certifikace
vybavení letadel dle CS-23.
Keywords: airworthiness certification, aircraft instruments, hardware, software.
1 Certification — the term and its definition by
a regulation system
Rules and processes of certification (airworthiness assurance for flight
operations) of aircraft and other aeroplane technology products are in
Europe generally formulated by EASA Part 21 regulation [1]. A particular certification for normal, utility, acrobatic and commuter category airplanes is CS-23 specification. The basic documents in EASA
document system are supplemented by other specifications, which are
connected with a qualification of an organization for aircraft and its
parts maintenance (Part 145), certifying of an aircraft maintenance
personnel (Part 66) and with an approval of an organization for maintenance personnel training (Part 147). The system of certification specifications comprises Acceptable Means of Compliance (AMC) and
the obliged European Technical Standard Orders (CS-ETSO) too.
The document system described in this way is not complete. It is
described in [2] in more detail with only reservation that the EASA
takes over gradually the JAA system and the referenced work does
not catch the evolution according to its time of publication. Current
state of documents is presented by EASA on its web site [3]. CAA CZ
has published on its site http://www.caa.cz/ in Czech language ”the
Processes“ based on EASA documents. References to some more
important of them are put to Bibliography clause points [3] to [6].
The way other regulation systems documents are presented is interesting from our point of view. So the standards of environmental tests
RTCA DO-160D (EUROCAE ED-14D) [7] and software standard
RTCA DO-178B (EUROCAE ED-12B) [8] are referenced in
CS-ETSO, Subpart A and in particular ETSO documents. In addition
to that the last document is a subject of AMC 20-115B. Other documents are not stated in CS-ETSO, Subpart A, but there is the possibility of including them in particular ETSO and addresses of seven
standardisation organisations are given without other details.
Nevertheless the unpleasant state lasts, when a low number of
AMC documents (they were preceded by similar ACJ documents in
JAA system) is in silence completed with AC documents from FAA
system. So meaningful AMC 25.1309 could be found for aircraft equipment in CS-25 Book 2, but corresponding AMC is not presented in
CS-23 and there is not (with exception of clauses 23.55 Acceleratestop distance and 23.75 Landing in Book 2) any reference to FAA system's AC23.1309 use. The status does not appear satisfying and it can
only be awaited that it will be in the future a subject of criticism and
then corrected by EASA. Some positive exception is a reference in
Book 2 to AC 23.1311-1 for instrument systems with electronic displays. Not only AC23.1309, but the documents SAE ARP4754, SAE
ARP4761, RTCA DO-178B, RTCA DO-254 and others too are referenced in AC23.1311-1.
2 Complex systems — particular requirements
A fundamental change of certification procedures arises with supplementing of instrument panel of aircraft by electronic system with data
presentation on displays. The function of particular subsystems cannot be assured by classical measuring procedures, the subsystems are
not sufficiently separated. Electromechanical systems of cockpit
instruments are replaced in permanently growing extent by electronic
and computing systems with program control.
The systems are supposed to be complex. They neednot be in some
cases developed specially as systems for avionics. They can be commercial of the shelf systems, in acronym known as COTS.
Standardisation organisations developed for certification of complex systems support a document system with complex coverage of
the problems. When we start from presentation published in [9], we
receive an extended document system in Figure 1.
In centre of the system are above-mentioned documents SAE ARP
4754 / ED-79, SAE ARP 4761, RTCA DO-178 / ED-12B, RTCA
DO-254 / ED-80. The documents are appended by other documents
52
●
●
●
●
Figure 1: EUROCAE and SAE document system for aircraft certification support
RTCA DO-160, DO-264, DO-278, DO-200A / ED-76 (Standards for
Processing Aeronautical Data) & DO-201A. Life cycle processes of
complex electronic hardware (containing e.g. PLD, FPGA or ASICs)
with co-operating software — called embedded systems, are covered
by such composed set of documents. In many cases they are systems,
whose controlling kernel is processor with a memory.
Interrelations in the system were described in article [10]. Primary
document of the system is ARP 4754 standard. It describes development processes of a system and processes of safety assessment in
interrelation and with links to different levels of functional hierarchy,
from aircraft level functions, through allocation to systems, systems
development and requirements allocation to hardware and software
objects, to actual system implementation. Other topic of the document
are support processes of systems and objects development related to
hardware and software life cycle, and certification co-ordination on
aircraft function level as one of support processes. Figure 2 reproduces the processes and interfaces according to [11].
The development procedure functions comprise generally:
●
●
●
system architecture design and allocation of requirements to
objects
design and construction of hardware and software
hardware and software integration
system integration
There are elaborated procedures in frame of system engineering for
some of the activities. System engineering, called sometimes as requirement engineering [12], is concentrated on requirements, which it
considers as objects, in the course of the whole procedure of system
development. It offers for the role a variety of procedures, by which
the requirements are identified, analysed, validated, tracked and controlled. Using of requirement engineering during the development of
aircraft systems linked with structure of relation of development process and safety assessment process in ARP 4754 is described by [13].
Figure 3 is taken from the paper, offering concept of product and
associated support products breakdown.
External higher level requirements (airline needs, other stakeholder
requirements) generate assigned specified requirements from aircraft
design to partial object design. Dependability requirement of aeroplane categories according to FAR Part 23 and Part 25 is described in
detail by Hlinka [14]. The same author described in [15] his practical
experience with reliability assessment using tests evaluation performed on small aeroplane and he summarised other results gained
during his work on the problems in article [16]. The works [17] and
[18] take a broader view of certification.
identification of aircraft level function, functional requirements
and functional interfaces
determination of functional failures consequences
allocation of functions to systems and persons
Figure 2: Relation of supporting processes to aircraft HW and SW lifecycle
Figure 3: Operational functions and associated life-cycle process functions breakdown
Documents denoted ED are created by EUROCAE (The European
Organisation for Civil Aviation Equipment). The standard ED-12B has
been used for years in software certification, however use of ED-80
for hardware is a matter of the last five years. Both the documents
define development assurance levels. So ED-80 defines five System
Development Assurance Levels; it describes for the two highest
(Catastrophic — A, Hazardous/Severe-major — B) levels additional
design assurance activities:
●
architectural mitigation techniques — dissimilar implementation, redundancy, monitors, isolation, partitioning, etc.
53
●
●
product service experience
advanced verification methods — elemental analysis, safety
specific analysis, formal methods
Criticality level is assigned to the developed software in ED-12B
according to software contribution to possible failure states, determined during system safety assessment procedure. The compliance of
the developed software with applicable requirements shall be demonstrated during airborne software certification testing. The cogency of
used testing method, usually of used software tool, shall be demonstrated too. The method of managing particular problems following
from ED-12B requirements is described in a range of documents published by Certification Authorities Software Team [19]. Table 1 shows
a merging of ED-12 [8] and ED-80 [20] classification definitions for
particular development assurance levels.
Description 4:
Description 5:
3/2005
impairing flight crew efficiency, or discomfort to occupants,
possibly including injuries.
Failure conditions that would not significantly reduce aircraft
safety, and which would involve flight crew actions that are
well within their capabilities. Minor failure conditions may
include: a slight reduction in safety margins or functional
capabilities, a slight increase in flight crew workload, such as
routine flight plan changes, or some inconvenience to occupants.
Failure conditions that do not affect the operational capability of the aircraft or increase flight crew workload
3 Reliability and safety in life cycle phases
The failure rate decreases at hardware during initial phase of life
cycle, then it stays constant in predominant part of product life and
the failure intensity growths to the life cycle end. Graphical representation is a bathtub curve, which is often under criticism but is permanently used. Initial failures (called quaTable 1 System Development Assurance Levels and their Definitions for HW and SW
lity failures too) satisfy to child mortality period, stress failures meet constant
failure rate period and the final part of
the curve is created by wear-out failures. The bathtub curve results by superposition of the partial curves.
The constant failure rate is an exponential distribution parameter. Initial
failures and wear-out failures can be
modelled by Weibull distribution with
appropriate coefficients.
EUROCAE ED-80 [20] document is
devoted to hardware development assurance and hardware life cycle. One of
the authors of the paper presented the
contents and importance of the document in frame of lecture cycle TEMPUS [21].
Particularities of software life cycle
lead to the following processes:
●
process of sw planning (it
defines and co-ordinates sw development activities and process of sw integration into project);
●
processes of sw development
(creating of sw product):
●
sw requirements process;
-
DAL ... Development Assurance Level
SCL ... Software Criticality Level
Description 1: Failure conditions that would prevent continued safe flight
and landing
Description 2: Failure conditions that would reduce the capability of the
aircraft or the ability of the flight crew to cope with adverse
operating conditions to the extent that there would be: a large
reduction in safety margins or functional capabilities, physical
distress or higher workload such that the flight crew could
not be relied on to perform their tasks accurately or completely, or adverse effects on occupants including serious or
potentially fatal injuries to a small number of those occupants.
Description 3: Failure conditions that would reduce the capability of the
aircraft or the ability of the flight crew to cope with adverse
operating conditions to the extent that there would be: a significant reduction in safety margins or functional capabilities,
a significant increase in flight crew workload or in conditions
sw design process;
sw code creation process;
integration process;
●
integral processes (they assure correctness, control and confidence
in sw life cycle processes and their outputs, they proceed concurrently with sw development):
-
sw verification process;
-
sw configuration management process;
-
sw quality assurance process;
-
certification agency liaison process.
There are one or more software life cycles defined in a project
through activity choice for every process, assignment of the activities
sequence and mapping the responsibility for the activities [8]. The
processes sequence is determined by process characteristics, which
are:
●
functionality and complexity;
●
volume of the sw;
54
●
●
●
●
requirements stability;
usability of preceding development results;
development strategy;
hw accessibility.
The software verification process comprises among others the software testing process. The process has two complementary goals:
●
●
demonstration of compliancy to requirements;
demonstration with high degree of confidence, that the errors
were eliminated, which could lead to unacceptable failure states
determined by system safety evaluation process.
4 Methods and means of assurance
and subsystems) and their assessment comes under ”Interfacing
Requirement Management Tools“ region described in [26]. To the
tools family belongs ”Requirements Extraction Tool“ [27] with some
properties close to ISVI created by Aeronautical Research and Test
Institute [28].
Our goal will be creation of a tool universal enough to be used for
certification of products from other industries as well, while achieving
the most effective operation possible and a user friendly environment
by incorporating in it some specialised modules. For aircraft industry
the selected region is certification of equipment according to CS-23
specification. One of the possible views of the concept GUI for VIS
(Verification Information System) is shown in Figure 5.
There is the whole line of procedures of software requirements validation and verification assurance methods. They are additionally different for systems differing in consequence of their functional diversity and other differences, e.g. location in aircraft.
One of the strongest aids used in this field is system engineering,
sometimes called requirement engineering, and its methods. Main
roles of requirement engineering are:
●
●
to address requirements as objects all along a system development process;
to provide a set of processes to identify, analyse, validate, trace,
manage requirements.
V-diagram in Figure 4 comprehends requirements validation and system or equipment development and production verification. It is simplified from detail figure in [22] for better insight.
Figure 5: GUI concept of Verification Information System
5 Summary
Figure 4: V-Diagram
There are various means supporting the requirement engineering methods. One very useful of them are formal languages and software
tools.
4.1 Formal languages
Formal methods called co-design or co-synthesis are used for modelling and design of complex hardware systems co-operating with software, which are usually categorised as embedded systems. There are
a variety of procedures, but they are generally based on four concepts:
—
concurrency;
—
hierarchy;
—
communication;
—
synchronisation
Some of them are very smart and risen to sw implementation, like e.g.
hybrid system model with structurally oriented architecture reliability
block diagram model with components modelled as finite automata
[23]. We recommend for more detail information concerning modelled hw/sw structures and for comparing of system specification languages the surveys published in [24] and [25]. In SAE standardised
AADL (Architectural Analysis&Design Language) cannot be marginalised from all the languages used in the field.
4.2 Tools
The necessary tool for requirement documents, analysis results and
test results organisation (in relation to particular aircraft equipment
The result of CS-23 aircraft certification rules and their implementation in the course of aircraft design and production is the finding that
they present enormous demands on both designer and certification
authority. In the contribution we have tried to take into account most
of the problems involved, from regulation system assessment through
the system of associated standards and the procedures emerging thereof, to the fact that certification of a complex aircraft system has to
be based on the use of a software tool. The tool concept suggested is
outlined in the end. The suggested VIS concept is determined for
making certification effortless and transparent for both sides of certification process.
References:
[1]
[2]
[3]
[4]
[5]
[6]
[7]
[8]
[9]
EASA Part 21: Certification of aircraft and related products, parts
and appliances, and of design and production organisations. In:
Official Journal of the European Union L243, 27.9.2003
Svoboda, M.: Požadavky evropských leteckých předpisů na systém
zabezpečení vývoje letecké techniky a jejich vztah k příslušným
požadavkům normy ISO 9001, VZLÚ, 2000
CAA-ST-076-0/04 Směrnice Postupy osvědčování vybavení; ÚCL
ČR, vydáno 16.7.2004
CAA-ST-075-0/04 Postupy pro získání oprávnění organizace
k výrobě (POA) pro výrobky, letadlové části a zařízení dle přílohy
nařízení Komise (ES) č. 1702/2003 (Část 21); ÚCL ČR, schváleno
21.5.2004
CAA-ST-081-0/04 Postupy pro vystavování osvědčení letové způsobilosti pro nově vyrobená sériová letadla v ČR organizacemi
oprávněnými k výrobě podle Části 21, Hlavy G; ÚCL ČR, vydáno
17.12.2004
CAA-ST-069-2/04 Postupy opravňování ETSO (pro žadatele);
ÚCL ČR, vydáno 31.3.2004, změna č. 2 z 24.6.2005
DO-160D Environmental Conditions and Test Procedures for Airborne Eqipment, Change 3, RTCA, December 2002
DO-178B Software considerations in Airborne Systems and Equipment Certification, RTCA, SC-167, December 1992
Forsberg, H.: Certification Issues in Avionics - COTS and Complex
Electronic Hardware; SAAB Technologies, 2004;
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Merkl, M., Nová, M.: Zdokonalování metodických nástrojů pro
zabezpečení spolehlivosti a bezpečnosti letecké techniky; Letecký
zpravodaj 2/2001, str. 56-58
ARP 4754 Certification Considerations for Highly-Integrated or
Complex Aircraft Systems; SAE SIRT Group; Apr 1996
NAS System Engineering Manual Version 3.0 09/30/04
Fraboulet, G., deChazelles, P.: Use of Requirement Engineering Discipline in support of A/C Engineering; Seditec, Airbus;
www.estec.esa.nl/conferences/02C24/2_dechazelles_fraboulet.pdf
Hlinka, J.: Aircraft Dependability Requirements Contained in Federal Aviation Regulations Part 23 and Part 25, Letecký zpravodaj
4/2002, pp.22- 25
Hlinka, J.: Dependability Assessment Possibilities for General Aviation Class Aircraft during Development and Certification; Letecký
zpravodaj 3/2003, pp. 48-51
Hlinka, J.: Application of Reliability Analysis During Certification of
GA Airplanes; Czech Aerospace Proceedings 3/2004, pp. 19-22
Mžik, T., Hlinka, J.: Reliability of aircraft in design and certification
stage. in: Proceedings of the second international conference "Reliability, Safety and Diagnostics of Transport Structures and Means
2005", Pardubice, 7-8 July 2005; pp. 239-247
Hlinka, J., Novák, J.: Reliability Analysis of Small Sport Airplane. in:
Proceedings of the second international conference "Reliability,
Safety and Diagnostics of Transport Structures and Means 2005",
Pardubice, 7-8 July 2005; pp. 119-127
Certification Authorities Software Team (CAST); Position Papers;
www.faa.gov/certification/aircraft/av-info/software/CAST.htm
ED-80 Design Assurance Guidance for Airborne Electronic Hardware, EUROCAE, 2000
Merkl, M.: Problematika bezpečnosti a JAR-23; Studijní materiály
[22]
[23]
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ke kurzu Tempus Phare; VTÚL a PVO Kbely, 31.10.2002
Tews, H.J.: TESTNET, A Network of excellence, Integrating of Testing Methodologies; http://www-lor.int-evry.fr/testnet/slides/TestNnet_kick_off_short.pdf
Rehage, D., Carl, U. B., Vahl, A.: Redundancy management of fault
tolerant aircraft system architectures - reliability synthesis and analysis of degraded system states; Aerospace Science and Technology
9 (2005); pp.337-347
Jerraya, A. A. et all: Multilanguage Specification for System Design
and Codesign; in System-level Synthesis, NATO ASI 1998, Kluwer
1999; http://tima.imag.fr/publications/files/rr/mss_54.pdf
Panagopoulos, I., Papakonstantinou, G., Alexandridis, N.: A Comparative Evaluation of Models and Specification Languages for
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http://skyblue.csd.auth.gr/~bci1/Panhellenic/243Panagopoulos.pdf
Jones, D. A. et all: Interfacing Requirements Management Tools In
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Seventh International Symposium of the INCOSE — Volume II,
August 1997; www.itmweb.com/essay544.htm
XTie-RT Requirements Tracer Software; Teledyne Brown Engineering,
http://scholar.lib.vt.edu/theses/available/etd-1598-132027/unrestricted/ch5.pdf
Merkl, M.: Začlenění textových informací do elektronické kanceláře
konstruktéra; Letecký zpravodaj 3/2000, str. 44-48
Czech Aerospace Research Centre — Workshop 2005: / Seminář CLKV 2005
Addendum:
Directory of Papers not Published in this Issue
Seznam nepublikovaných příspěvků
To bring a complete overview of this year's proceedings the following list sums up all other
papers presented at the 2005 conference that
could not be published in this issue due to capacity restriction.
Pavlica, R. / VUT:
New Technologies for Advanced Composite Structures
Nové technologie pro moderní kompozitní konstrukce
Tůma, J. / VZLÚ:
Aerodynamic Optimization of a Compressor Blade
Aerodynamická optimalizace lopatky kompresoru
Poul, R. / ČVUT:
Mass Optimization of UL Plane Cold Propulsor Drive
Shaft
Hmotnostní optimalizace transmisního hřídele pohonu UL
letounu se studeným propulsorem
Chudý, P. / VUT:
Safety Analysis of Aircraft System
Bezpečnostní analýza letadlových systémů
Weigel, K. / ČVUT:
Frequency Tests on TL-32 Typhoon
Ověřovací frekvenční zkoušky na letounu TL-32 Typhoon
Mališ, M. / VUT:
Effect of Geometry on Glider Fuselage Resistance
Vliv geometrie na odolnost trupu kluzáku
Patočka, K., Jamróz, T. / VZLÚ:
Development of Formulas in Application of the Ritz Method to Composite Plate Buckling
Odvození vztahů pro ztrátu stability kompositových
desek za použití Ritzovy metody
Mihalides, D. / VUT
Effect of Matrix Systems on Composite Mechanical Properties
Vliv matricových systémů na mechanické vlastnosti kompositu
Urík, T. / VUT:
Fatigue Tests of Adhesive-Riveted Joints
Únavové zkoušky lepeno-nýtových spojů
Posters / Posterové prezentace
Nová, M. / VZLÚ:
Payload Integrity Analysis of a Scientific Satellite
Analýza bezporuchovosti užitečného zatížení vědecké družice
56
Valeš, M. / VZLÚ, Kachlík, P. / VUT:
Removing Paint from Composites by Dry-ice Blasting
Odstranění nátěrů z kompozitních materiálů revoluční
technologií-otryskávání suchým ledem CO2
Fedossov, V. / VZLÚ:
Securing Warranty in Space Research Projects
Zajištění garance v projektech kosmického výzkumu
Anderle, P. / ČVUT:
Experimental Facilities for Aerodynamics
Experimentální zařízení pro aerodynamická měření
Růžička, P. / ČVUT:
Static Tests of Thin-walled Shell Shafts — Checking Composite and Hybrid Characteristics
Statické zkoušky tenkostěnných skořepinových hřídelů —
ověřování charakteristik kompositních a hybridních konstrukcí
Room for Your Notes
Environment in Evektor EV-55 Aircraft
Colour illustrations to the article published on pages 9-10.
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ISSN 1211 - 877X
Introductory Lecture to the ARC 2005
Slovo úvodem k semináři CLKV 2005
in Aerodynamic Optimisation
Použití metody Nelder-Meada při
aerodynamické optimalizaci
of Landing Gear
Analýza životnosti podvozku na
základě výpočtů MKP
Composites — A New Trend in the
Airplane Control System
Impeller Design
Počítačové modelování mikroklimatu
v kabině dopravního letadla
Evektor EV-55
Kompositní materiály — nový trend
v konstrukci řízení letounů
Tlakové zkoušky pro stanovení
materiálových charakteristik použitých
v numerickém modelu FSW
Návrh oběžného kola odstředivého
kompresoru s tandémovým
Užití spektrální dekompozice při snižování
hluku vrtulových letadel
Úvod do problematiky recyklace
termosetických kompositních materiálů
Safety Analyses of Aircraft
Avionic Systems
Aerodynamic Design of V44 Model Propeller
and Aeroacoustic Characteristic Calculation
Rozbor bezpečnosti avionických systémů
letadla
Aerodynamický návrh modelové vrtule V44
Výpočet aeroakustických charakteristik
N o v e m b e r
Contents / Obsah
2 0 0 5
ISSN 1211—877X
CZECH
AEROSPACE
Proceedings
LETECK Ý
zpravodaj
In this issue:
Czech Aerospace
Research Centre
CLKV
Proceedings of the
5th Annual
Workshop held at
Holany, Czech
Republic
November 3 to 4,
2005
Zatížení vodorovných ocasních ploch
při manévru
Centrum leteckého
a kosmického
výzkumu
Použití neuronových sítí při analýze
plynové cesty turbinového motoru
CLKV
Měření distribuce tlaku profilu na pozemní
pojízdné laboratoři
Analytical Methods for the Calculation of
Buckling in Composite Sandwich Panels
Výpočet vzpěrné pevnosti kompositových
sendvičových desek
Reliability and Safety Points of View
Certifikace složitého vybavení z pohledu
zajištění spolehlivosti a bezpečnosti
© C Z E C H AE R O S PAC E
M A N U FAC T U R E R S A S S O C IAT I O N
Sborník vybraných
referátů
přednesených na
5. ročníku
semináře CLKV
Holany, 3. — 4.
listopadu 2005
No. 3 / 2005

czech aerospace - Výzkumný a zkušební letecký ústav

Transcription

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