Tran Sonic Air Foils

7/28/2019 Tran Sonic Air Foils

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Lesson 26Review: Transonic flowfields are inherently nonlinear Advances in both experimental and computational methods

were required and achieved

Today: Discussion of transonic airfoil characteristics and design goals

Primarily associated with

Richard Whitcomb, Feb 21, 1921 Oct. 13, 2009


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Subsonic Linear Theory, Even with the

Compressibility Correction,

Cant Predict Transonic Flow!From Desta Alemayhu


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A real example

Illustrates tricks used to get calculations to agree with test data


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REVIEW: Obtaining CFD solutions

Grid generation Flow solver

Typically solving 100,000s (or millions in 3D)of simultaneous nonlinearalgebraic equations

An iterative procedure is required, and its noteven guaranteed to converge!

Requires more attention and skill than lineartheory methods

Flow visualization to examine the results


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Airfoils

Mach number effects: NACA 0012

-1.5

-1.0

-0.5

0.0

0.5

1.0

1.50.0 0.2 0.4 0.6 0.8 1.0

CP(M=0.50)

CP(M=0.70)

CP(M=0.75)

Cp

X/C

NACA 0012 airfoil,FLO36 solution, = 2

M=0.50

M=0.70M=0.75


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Angle of attack effects: NACA 0012-1.50

-1.00

-0.50

0.00

0.50

1.00

1.50

0.00 0.20 0.40 0.60 0.80 1.00

CP( = 0)

CP( = 1)

CP( = 2)

Cp

x/c

FLO36NACA 0012 airfoil,M= 0.75


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Traditional NACA 6-series airfoil

-0.10

0.00

0.10

0.20

0.00 0.20 0.40 0.60 0.80 1.00

y/c

x/c

Note small leading edge radius

Note continuous curvatureall along the upper surface

Note low amount of aft camber

-1.50

-1.00

-0.50

0.00

0.50

1.00

1.500.00 0.20 0.40 0.60 0.80 1.00

Cp

x/c

FLO36 prediction (inviscid)

M= 0.72, = 0, CL

= 0.665

Note strong shock

Note that flow acceleratescontinuously into the shock

Note the low aft loading associatedwith absence of aft camber.


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A new airfoil concept - from Whitcomb

Progression of the Supercritical airfoil shapeNASA Supercritical Airfoils, by Charles D. Harris, NASA TP 2969, March 1990

1964

1966

1968


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What the supercritical concept achieved

From NASA Supercritical Airfoils, by Charles D. Harris, NASA TP 2969, March 1990

Section drag at CN

= 0.65 Force limit for onset of upper-surface boundary layer separation


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And the Pitching Moment



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How Supercritical Foils are Different



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Supercritical Airfoils

-0.10

-0.05

0.00

0.05

0.10

0.15

0.20

0.00 0.20 0.40 0.60 0.80 1.00

y/c

x/c

Note low curvatureall along the upper surface

Note large leading edge radius

Note large amount of aft camber

-1.50

-1.00

-0.50

0.00

0.50

1.00

1.500.00 0.20 0.40 0.60 0.80 1.00

x/c

Cp

FLO36 prediction (inviscid)

M= 0.73, = 0, CL

= 1.04

Note weak shockNote that the pressuredistribution is "filled out",providing much more lifteven though shock is weaker

Note the high aft loadingassociated with aft camber.

"Noisy" pressure distribution is associatedwith "noisy" ordinates, typical of NASAsupercritical ordinate values


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Whitcombs Four Design Guidelines

An off-design criteria: a well behaved sonicplateau at M= 0.025 below the design M

Gradient of pressure recovery gradual enough toavoid separation in part: a thick TE, say 0.7% on a 10/11% thick foil

Airfoil has aft camber so that design angle ofattack is about zero, upper surface not sloped aft

Gradually decreasing velocity in the supercriticalregion, resulting in a weak shock

Read NASA Supercritical Airfoils, by Charles D. Harris,

NASA TP 2969, March 1990, for the complete story


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Airfoils 31 and 33


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Off Design


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Foils 31 and 33 Drag


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NASA Airfoils Developed Using the Guidelines

fromNASA Supercritical Airfoils, by Charles D. Harris, NASA TP 2969, March 1990

Filled symbols denote

airfoils that were tested


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NASA Airfoil Catalog

Note: watch out for coordinates tabulated in NASA TP 2969!


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Frank Lynchs Pro/Con Chart

for supercritical airfoils

F.T. Lynch, Commercial TransportsAerodynamic Design for Cruise Performance

Efficiency, in Transonic Aerodynamics, ed. by D. Nixon, AIAA, 1982.


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Airfoil Limits: the Korn Eqn.

We have a rule of thumb to let us estimate whatperformance we can achieve before drag divergence By Dave Korn at NYU in the 70s

MDD

+C

L

10+

t

c

=

A

A= 0.87 for conventional airfoils (6 series)

A= 0.95 for supercritical airfoils

Note: the equation is sensitive to A

This is an approximation until CFD or WT results arrive!


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Airfoil Limits

Shevell and NASA Projections Comparedto the Korn Equation

0.65

0.70

0.75

0.80

0.85

0.90

0.04 0.06 0.08 0.10 0.12 0.14 0.16 0.18

t/c

MDD

Shevell advanced transonicairfoil estimate

Korn equation, A

= .95

Korn equation, A

= .87

Shevell estimate,mid 70's transportairfoil performance

0.65

0.70

0.75

0.80

0.85

0.90

0.02 0.06 0.10 0.14 0.18t/c

MDD

CL

0.4

0.7

1.0NASA projection

Korn equation estimate, A

= .95

In W.H. Mason, Analytic Models for Technology Integration in

Aircraft Design, AIAA Paper 90-3262, September 1990.


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For the curious: the airfoil used on the X-29


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Just when we thought

airfoil design was

finished

See Henne, Innovationwith Computational

Aerodynamics: The

Divergent Trailing Edge

Airfoil, inApplied

Computational

Aerodynamics, P.A.

Henne, ed., AIAA

Progress in Aero Series,

1990


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Used on the MD-11

resisted in Seattle!


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Take a Look at

the PressureDistribution

Comparison of the

DLBA 243 and the

DLBA 186 Calculated

Pressure Distributionat M = 0.74


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To Conclude:

You now know the basis for

transonic airfoils

Tran Sonic Air Foils

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