AERODYNAMICS TOOLS AND METHODS IN AIRCRAFT DESIGN … · • The BWB is an aircraft which...
Transcript of AERODYNAMICS TOOLS AND METHODS IN AIRCRAFT DESIGN … · • The BWB is an aircraft which...
A RAPID 3D AERODYNAMIC PREDICTIONMETHOD FOR BLENDED-WING-BODYCONCEPTS
Dr Davide Di Pasquale & Dr Simon Prince
AERODYNAMICS TOOLS AND
METHODS IN AIRCRAFT DESIGN
London, 14-15 Oct 2019.
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1) Motivation – rapid and “appropriate” prediction methods.
2) Brief overview of the method
3) Validation test cases
1. The ARA RBC12 Wing / Body Configuration
2. The Cranfield Eagle-Ray BWB aircraft Configuration
4) Graphical User Interface Development
5) Eagle-Ray BWB aircraft optimisation
6) Conclusions
Outline
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Conceptual Design Vision - Quicker, Cheaper and Right First Time
1. Motivation – rapid and “appropriate” prediction methods.
Source: Rolls-Royce Source: Sandy Monro (Ford Motor Company) – Lean Design Philosophy, 1988: Source: Sirirojvisuth, 2012
• Studies have shown that “product (early) design” has the greatest influence on productivity
improvement and downstream costs.
• Correcting the effects of “poor design” can be prohibitively expensive and have tangible
impact on market share and/or business performance.
• The aerodynamic design and architectural definition of a modern aircraft requires the ability
to rapidly and cost efficiently analyse and optimise a vehicle configuration.
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• Need to develop methods which are accurate enough but are quick
enough to produce the required design data to allow efficient design
analysis / trade-off studies.
• Such data must be of sufficient accuracy, in terms of overall and local lift
and drag forces, that the performance trends are correctly captured
• Benefits: 1) Much reduced cost of conceptual design.
2) Improved reaction time to market forces.
3) Ability to better drive out design “mistakes” early in
cheapest stage of the design process.
1. Motivation – rapid and “appropriate” prediction methods.
Conceptual Design Vision - Quicker, Cheaper and Right First Time
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• Viscous Full (Non-Linear) Potential Equations coupled with Integral Boundary
Layer equation solver (semi-inverse, swept / tapered wing integral boundary layer
method)*
• Assumptions: flow is steady
flow has no separation (laminar bubbles captured).
flow is irrotational.
flow is isentropic, (only weak shock waves).
• Boundary Layer Equations directly solved in terms of the required variables (δ,
δ*,θ, H ). ie: no costly and problematic post processing.
• The code allows the wing geometry to be input as a series of section profiles to be
defined from the root to the tip, along with the corresponding location of the local
leading edge and the chord length.
• Structured mesh (automatic generation from wing and body section data).
*Ashill, P. R. & Smith, P. D. “An integral method for calculating the effects on turbulent boundary layer development on sweep
and taper”, RAE Technical Report, TR83053. June 1983.
2. Brief Overview of the method
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3.1 The ARA RBC12 Wing / Body Configuration.
3. Validation test cases
• Quarter chord sweep of 25o, semi-span of 1.085m, mean aerodynamic
chord of 0.279m, and an aspect ratio of 7.78.
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3.1 The ARA RBC12 Wing / Body Configuration.
3. Validation test cases
• Comparison with NS / DDES
Mach 0.8, Rec = 3.75x106 condition.
• NS / DDS for 3 points: 15 days (~20 millioncells, 128 core processors)
• VFP Pitch Sweep: ~1 hour (135,000 cells, 1processor, standard laptop.)
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3.1 The ARA RBC12 Wing / Body Configuration.
3. Validation test cases
Mach = 0.8
a = 2.4o case
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3.1 The ARA RBC12 Configuration.
3. Validation test cases
M∞=0.8
= 3.76o,
Rec=3.75x106(DPSP)(Time Av.)
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3. Validation test cases
• The BWB is an aircraft which integrates the wing and the fuselage, and which does not
contain any tail for flight control.
• The BWB centre-body provides lift which improves the aerodynamic performance by
reducing the wing loading, compared to the cylindrical fuselage of a conventional
aircraft.
• The Blended-Wing Body (BWB), has the potential to reduce the fuel consumption by
ingesting the boundary layer (BLI) and lower the acoustic impact since the exhaust noise
can be shielded to a certain degree by the wing.
• The BWB is quiet, strong, and because of its economical performance is a promising
candidate for the future large airliner.
R. H. Liebeck. Design of the blended-wing-body subsonic transport. In AIAA Aerospace Sciences Meeting and Exhibit, 40th, Reno, NV;
UNITED STATES; 14-17 Jan. 2002. Reston, 14-17 Jan. 2002 2002. ISBN 0146-3705. URL http://www.vicomplex.hu/arep/BoeingBWB.pdf.
3.2 The Cranfield Eagle-Ray BWB Configuration
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3. The BW-11 Eagle-Ray Wing-Body
Airframe alone analysis (engines modelling comes later) without winglets.
VFP versus RANS comparison and then optimisation.
a)V
FP M
esh
b)
RA
NS
Mes
h
~15M cells
135,432 cells
15 constant spanwise stations on the BWB were extracted from
the CAD definition
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3. The BW-11 Eagle-Ray Wing-Body
a) Lift coefficient b) Drag coefficient
c) Lift to drag ratio
Comparison of predicted force characteristics, M=0.80, Rec=3.7x108.
VFP ~7 min for point (1 CPU)
Fluent ~5 days for point (16CPUs)
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4. VFP Graphical User Interface
• The need to speed up the manual optimisationprocess has led to working on the graphical userinterface development within Matlab, avoiding towork with strictly formatted ASCII files.
• Ability to rapidly and efficiently modify the wingdesign variables with the geometry module.
• Ability to perform a rapid and thoroughassessment of the simulation results with the post-processing module in order to optimise thegeometry.
With this toolset, the designer could therefore rapidly perform trade-off analyses to improve
their configuration designs and obtain good lift to drag characteristics with safe buffet onset
margins.
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4. VFP GUI Geometry Module
1. Import a file (.GEO, excel or .txt formats
accepted) containing the wing geometry.
2. Export the desired geometry in a GEO file
format.
3. Save results and keep record of the new
wing configuration.
4. Reset all the variables to the baseline case.
5. Select the 3d plot or the desired section plot.
6. Introduce the new wing design variables for
the desired section.
7. Compute the modifications for the selected
section.
8. Compute a global increment of twist or
dihedral for the selected sections (more than
one if it is desired).
9. Plot the twist and dihedral spanwise
distributions.
10. Improve controls panel is used to compute
smooth and fast changes on twist, dihedral
and leading edge x coordinate variables
following a parabolic function.
11. The modifications are displayed in the table
in order to keep record.
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Layout Overview
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4. VFP GUI Post-Processing Module
• The spanwise loading can be plotted
selecting the multiple options in the
highlighted pop up menu presented
below. The lift, drag and pitching
moment coefficients, as well as, the lift
coefficient normalized by the geometric
mean chord can be visualized.
it offers the ability to compare multiple results.
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4. VFP GUI Post-Processing Module
1. Select upper or lower surface display.
2. Select the desired simulation step
comparison.
3. Selection of pressure coefficient or Mach
number contour plot.
4. Enable or disable iso lines.
5. Enter the desired contour lines.
6. Apply threshold value to assess the critical
zones.
7. Save results.
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5. Blended Wing Body: Optimisation
Testing Method
• Modify Twist and Dihedral.
• Constant Planform.
• Flow Conditions – Re= 9·106 based on
MAC chord with cruise at 35kft.
• Optimised design, Mach 0.80 AoA 0.0.
• Tested Mach range of 0.75 up until VFP
failure due to flow separation.
• Buffet-Onset Curve produced. BWB Mesh Visualisation inside the GUI
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5. Blended Wing Body: Optimisation
Key Points of Baseline Testing
• Separation at M 0.75 due to Upper SurfaceLeading Edge shock induced separation.
• Maximum CL limited by outer wing shockformation, leading to separation.
• High speed separation caused by adversepressure gradients on the Lower SurfaceTrailing Edge.
• Spanwise Loading showed a very undesirabledistribution.
BWB Baseline Buffet Curve
CP and CF Graph showing outer wing
Shock, Mach 0.85 AoA 1.8 Eta=0.925.
Section Twist -0.92̊
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5. Blended Wing Body: Optimisation
Twist Treatment
• Changes to twist compared at Mach 0.80,
Alpha 0.0.
• Aim to create an elliptical/triangular loading
distribution.
• Multiple iterations results in final twist design.
• At comparable flow conditions, CL increased
with increased CL/CD
CL CD CL/CD
Baseline 0.052 0.0085 6.13
BWB-T15 0.080 0.0088 9.09
Flow Conditions: Mach 0.80, AoA 0.0
Baseline vs BWB-T15 Spanwise
Loading Comparison
Comparison of Twist Distributions
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5. Blended Wing Body: Optimisation
Dihedral Treatment
• Increasing dihedral seemed to increase CL
• Large proportion of increased lift coming from inboard
wing
CL CD CL/CD
Baseline 0.052 0.0085 6.13
BWB-D18 0.056 0.0085 6.59Comparison of Dihedral Distributions
Flow Conditions: Mach 0.80, AoA 0.0
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5. Blended Wing Body: Optimisation
Final Results
• Maximum CL increased at all but one flow speed.
• CL increased at comparable flow conditions
• Maximum Alpha increased at lower Mach numbers
• Maximum Airspeed increased.
• Increased twist on inboard sections delayed lower surface flow separation.
• At Mach 0.89, optimised design performs worse, however separation location is more favourable.
• Design point of CL=0.152 and Mach 0.80 chosen.
• MDD occurs at between M 0.85-0.87.
Comparison of BWB vs Optimised Buffet
Curves
Optimised MDD at CL=0.152
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• No need to go to highest fidelity methods ….. “because they are there”.
• Much better (balancing cost, time to solution and accuracy) to go forappropriate fidelity methods.
• Using the FP equations, coupled with the turbulent integral boundarylayer equations, has demonstrated both the accuracy and the efficiencyof the method for attached flow cases, prior to buffet onset, which arerelevant to the transonic cruise condition.
• This method shows good applicability in the early design stages ofaircraft design, and has ultimately shown that can be used as a rapidoptimisation method.
• It can be the basis for an automatic and multi-variables optimisation tofurther optimise the design (Work in progress!!!)
6. Conclusions.
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[1] Smith, P. D., “A calculation method for the turbulent boundary layer on an infinite yawedwing in compressible, adiabatic flow”, ARC CP 1268. 1974.
[2] Full-potential (FP) method for three-dimensional wings and wing-body combinations –inviscid flow. Part I: Principles and results. ESDU 02013, June 2002 (with Amendment A, May2006).
[3] Viscous full-potential (VFP) method for three-dimensional wings and wing-bodycombinations. Part 1: Validation of VFP results with experiment and comparisons with othermethods. ESDU 13013.
[4] Von Karman, T. “Calculation of pressure distribution on airship hulls” NACA TM 574, 1930.
[5] De Jarnette, F. R., Ford, C. P. & Young, D. E. “A New Method for Calculating SurfacePressures on Bodies at an Angle of Attack in Supersonic Flow” AIAA Paper 79-1552. AIAA12th Fluid & Plasma Dynamics Conference, Williamsburg, VA, July 1979. doi:10.2514/6.1979-1552
[6] Ashill, P. R. & Smith, P. D. “An integral method for calculating the effects on turbulentboundary layer development on sweep and taper”, RAE Technical Report, TR83053. June1983.
[7] Liebeck R. H. Design of the blended-wing-body subsonic transport. In AIAA AerospaceSciences Meeting and Exhibit, 40th, Reno, NV; UNITED STATES; 14-17 Jan. 2002. Reston,14-17 Jan. 2002 2002. ISBN 0146-3705.
[8] Polhamus, E. C. Application of the Leading-Edge Suction Analogy of Vortex Lift to theDrag Due to Lift of Sharp-Edge Delta Wings. Langley, Washington D.C: NASA, 1968.
References