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.

[email protected]

[email protected]

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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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• 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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Mach = 0.8

a = 2.4o case

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3.1 The ARA RBC12 Configuration.


M∞=0.8

= 3.76o,

Rec=3.75x106(DPSP)(Time Av.)

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• 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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1 2 3 4 5 6

7

10

8 9

11

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.

11

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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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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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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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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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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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

Thanks for your attention

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