Laminar Flows 04

8/6/2019 Laminar Flows 04

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Laminar Flow

Rodney Bajnath, Beverly Beasley, Mike Cavanaugh

AOE 4124March 29, 2004


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Introduction

Why laminar flow?

Less skin friction Lower drag

Skin Friction: Laminar vs. Turbulent

0

0.002

0.004

0.006

0.008

0.01

1.0E+05 1.0E+06 1.0E+07

Reynolds Number

Laminar

Turbulent


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Natural Laminar Flow NACA 6-Series Airfoils

Developed by conformal transformations, 30 50% laminar flow

Advantages: Low drag over small operatingrange, high Clmax

Disadvantages: Poor stall characteristics,susceptible to roughness, high pitch moment,very thin near TE

Drag bucket: pressure distributions causetransition to move forward suddenly at end oflow-drag Cl range

Minimum pressure at transition location

NACA Report No. 824


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Natural Laminar Flow

NACA 6A-Series

30 - 50% laminar flow

Eliminated TE cuspEssentially same lift and

drag characteristics as 6-

series

NACA Report No. 903


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Comparison of NACA 6- and 6A-Series PressureDistributions

-1

-0.5

0

0.5

0 0.2 0.4 0.6 0.8 1

x/c

NACA 64-012

NACA 64-012A

Natural Laminar Flow

NACA 64-012: xtrupper = 0.5932, xtrlower = 0.5932

NACA 64-012A: xtrupper = 0.6214, xtrlower = 0.6215

XFOIL


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Natural Laminar Flow NLF Airfoils

Aft-loaded airfoils with cusp at TE (Wortmann or Eppler sailplane airfoils) Front-loaded airfoil sections with low pitching moments (Roncz-developed

used on Rutan designs or canards)

Also NASA NLF- and HSNLF-series, DU-, FX-, and HQ- airfoils

Inverse airfoil design based on desired pressure distribution, capitalize onavailability of composites

Low speed and high speed applications

Codes used for design include Eppler/Somers and PROFOIL

Up to 65% laminar flow

Drag as low as 30 counts

1. NASA Contractor Report No. 201686, 1997.

2. Lutz, Airfoil Design and Optimization, 2000.

3. Garrison, Shape of Wings to Come,Flying1984.

4. NASA Technical Memorandum 85788, 1984.


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Natural Laminar Flow: Case Study

SHM-1 Airfoil for the Honda Jet

Lightweight business jet, airfoil inversely designed, testedin low-speed and transonic wind tunnels, and flight tested

Designed to exactly match HJ requirements High drag-divergence Mach number

Small nose-down pitching moment

Low drag for high cruise efficiency

High Clmax

Docile stall characteristics

Insensitivity to LE contamination

Fujino et al, Natural-Laminar-

Flow Airfoil Development forthe Honda Jet.


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(Continued)

Requirements

Clmax = 1.6 for Re = 4.8x106, M = 0.134

Loss of Cl less than 7% due to contamination

Cm > -0.04 at Cl = 0.38, Re = 7.93x106, M = 0.7

Airfoil thickness = 15%

MDD > 0.70 at Cl = 0.38

Low drag at cruise


Flow Airfoil Development forthe Honda Jet.


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(Continued)

Design MethodEppler Airfoil Design and Analysis Code

Conformal mapping, each section designedindependently for different conditions

MCARF and MSES Codes Analyzed and modified airfoil

Improved Clmax and high speed characteristics Transition-location study

Shock formation

Drag divergence Fujino et al, Natural-Laminar-Flow Airfoil Development for

the Honda Jet.


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(Continued) Specifications:

Clmax = 1.66 for Re = 4.8x106, M = 0.134

5.6% loss in Clmax due to LE contamination (WT)

Cm = -0.03 at Cl = 0.2, Re = 16.7x106

(Flight) Cm = -0.025 at Cl = 0.4, Re = 8x10

6 (TWT)

MDD = 0.718 at Cl = 0.30 (TWT)

MDD = 0.707 at Cl = 0.40 (TWT)

Cd = 0.0051 at Cl = 0.26, Re = 13.2x106

(TWT) Cd = 0.0049 at Cl = 0.35, Re = 10.3x10

6 (WT)


Flow Airfoil Development for

the Honda Jet.


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Laminar Flow Control

stabilize laminar boundary using distributed suction through a perforated

surface or thin transverse slots

plenum chamber

outer skin

inner skin

Boundary layer thins and becomes fuller across slot

Benefits

A laminar b.l. has a lower skin friction coefficient (and thus lower drag)

A thin b.l. delays separation and allows a higher CLmaxto be achieved

Ref: McCormick, Aerodynamics, Aeronautics and Flight Mechanics, pg. 202.


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Notable Laminar Flow Control Flight Test Programs

Date Aircraft Test Configuration LF Result Comments

1940 Douglas B-18

(NACA)

2-engine prop

bomber

NACA 35-215

10x17 wing glove section

suction slots first 45% chord

LF to 45% chord

(LF to min Cp)

RC = 30x106

Engine/prop noise

effected LFsurface quality issues

1955 Vampire

(RAE)

single engine jet

upper surface wing glove

suction - porous surface

full chord suction

full chord LF

M~0.7 / RC=30x106

Monel/Nylon cloth

0.007 perforations

1954-

1957

F-94(Northrup/USAF)

jet fighter

NACA 63-213

upper surface wing glove

suction 12, 69, 81 slots

Full chord LF

0.6 < M < 0.7

RC = 36x106

at Mlocal >1.09 shocks

caused loss of LF

1963-1965

X-21(Northrup/USAF)

jet bomber

30 sweep

new LF wings for program

suction through nearly full spanslots both wings

full chord LF

RC = 47x106

effects of sweep on LFencountered

1985-1986

JetStar

(NASA)

4-engine business jet

two leading edge gloves

Lockheed slot suction & liquidleading edge protection

McDD perforated skin & andbug deflector

LF maintained to frontspar through two years

of simulated airlineservice

no special maintenancerequired lost LF in

clouds & during icing

LE protection effective

Ref: Applied Aerodynamic Drag Reduction Short Course Notes, Williamsburg,VA 1990.


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Why Does LFC Reduces Drag?

removes turbulent boundary layer

XFOIL output


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Why Does LFC Increases CLMAX?

move boundary layer separation point aft

x - ft

-1.0 -0.8 -0.6 -0.4 -0.2 0.0 0.2 0.4 0.6 0.8 1.0

Cp

0.0

0.2

0.4

0.6

0.8

1.0

-1.0(2196)

-0.25(759)

-0.0625(276) -0.015625(108)

m = 1/4

x0

= 1.0 ft

x0

= 0.25 ft

x0

= 0.0625 ft

x0

= 0.015625 ftReynolds Number = 6x10

6

Ref: A.M.O. Smith, High Lift Aerodynamics, Journal of Aircraft, Vol. 12, No. 6, June 1975


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Raspet Flight Research Laboratory Powered Lift Aircraft

Piper L-21 Super Cub (1954)

distributed suction - perforated skins

CLMAX= 2.16 4.0

2.0 Hp required for suction(Ref: Joseph Cornish, A Summary of the Present State of the Art inLow Speed Aerodynamics, MSU Aerophysics Dept., 1963.)

Cessna L-19 Birddog (1956)

distributed suction - perforated skins

CLMAX= 2.5 5.0

7.0 Hp required for suction(Ref: Joseph Cornish, A Summary of the Present State of the Art inLow Speed Aerodynamics, MSU Aerophysics Dept., 1963.)

Photographs Courtesy of the Raspet Flight Research Laboratory


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Suction Power Required for 23012 Cruise Condition

leading edge x (ft) trailing edge

0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0

-0.4

-0.3

-0.2

-0.1

0.0

NACA 23012

cruise CL = 0.410,000 ft.180 kts (303.6 ft/s)

adverse pressure gradient

dx

dUv

e

w= 18.2

Suction velocity required to maintainincipient separation of the laminar b.l

and prevent flow reversal is given by:

Joseph Schetz, Boundary Layer Analysis, Equation (2-37)

0.035

0.0025 dia

45 chord

12span45 x 12 grid 439,470 holes

Preq = .00318 Hp / foot of span*

*assumes:use highest vw and p in calculation

discharge coefficient of 0.5pump efficiency of 60%


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Laminar Flow Control Approaches

leading edge x (ft) trailing edge

0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0

-0.4

-0.3

-0.2

-0.1

0.0

NACA 23012cruise C

L= 0.4

10,000 ft.180 kts (303.6 ft/s)

adverse pressure gradient

1). Leading Edge Protection

2). Distributed Suction (perforated skin or slots)

3). Hybrid Laminar Flow ControlRef: Applied Aerodynamic Drag Reduction Short Course Notes,.Williamsburg,VA 1990.


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Laminar Flow Control Problems/Obstacles

Sweep

Attachment line contamination (fuselage boundary layer)

Crossflow instabilities (boundary layer crossflow vortices)

Manufacturing tolerances / structure

Steps, gaps, waviness

Structural deformations in flight

System complexity

Ducting and plenums

Hole quantity and individual hole finish

Surface contamination Bypass transition (3-D roughness)

Insects, dirt, erosion, rain, ice crystals

Ref: Applied Aerodynamic Drag Reduction Short Course Notes, Williamsburg,VA 1990.

Ref: Mark Drela, XFOIL 6.9 User Guide, MIT Aero & Astro, 2001


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Boundary Layer Transition Flight Tests on GlasAir

Oil flow tests on GlasAir (N189WB)

Raspet Flight Research Laboratory

August 1995

200 KIAS

5500 ft pressure altitude

Airfoil: LS(1)-0413mod GAW(2)

Mean aerodynamic chord: 44.1 in.

Re 7.5x106

Cruise CL 0.2


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Drag Benefit of Laminar Flow


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CENTURIA

4 Passenger Single Jet Engine GA Aircraft

CompetitionCirrus SR22Cessna 182

Targets existing General Aviation pilots

Cost ~ $750,000

International Senior Design ProjectVirginia Tech and Loughborough University


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Centuria Design Details Cruise altitude 10,000ft

Cruise Speed 185kts

Range 770nm

Take-off run 1575ft Aspect Ratio 9.0

Wing Area 12.3m2/132.39ft2

Thrust 2.877kN/647lbs

MTOW 1360kg/2998lb

Fuel Volume 773 litres/194 USG

Stall Speed 68kts (Clean) 55kts (Flap)


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Drawing byAnne Ocheltree & Nick Smalley


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Calculating Laminar Flow60%

100%Laminar Turbulent

Laminar Turbulent

Wing & Tail

Fuselage

0005.0Re

328.1

0032.0)144.01(Re)(log

455.065.0258.2

10

==

=+

=

40% 100%


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Fuselage Laminar to max thickness

Wing60% LM flow upper and lower surface

V-Tail60% LM flow upper and lower surface

Structure SWET (in2

) Turb C d Lam C d % Reduction

Wing 224.89 0.00875 0.00268 69.41

Tail 58.39 0.00211 0.00070 67.05

Fuselage 295.87 0.00975 0.00473 51.51

SREF (in2

) 132.72

Mcruise 0.29

Recruise 5.88E+06


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Reduction in Drag from Laminar flow

0

0.005

0.01

0.015

0.02

0.025

Turb Cd Lam Cd

Cd

Fuselage

Tail

Wing


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Centuria NLF Manufacturing Tolerances

Rh,crit hcrit (in.)

900 0.0072 inches

1800 0.0143 inches

2700 0.0215 inches

15,000 0.1195 inches

Carmichaels waviness 0.0139 inch/inchcriteria

Ref: A.L. Braslow, Applied Aspects of Laminar-Flow Technology, AIAA 1990

h


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Conclusions

Natural Laminar Flow Improvement of materials and computational methods allows

inverse airfoil design for desired characteristics or specific

configurations

Laminar Flow Control LFC is a mature technology that has yet to become

commercially viable

Drag Benefit on Centuria

61% reduction in skin friction drag due use of laminar flowon wings, tail and fuselage


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ReferencesAbbott, I.,H., Von Doenhoff, A.,E., Stivers, L.,S., Summary of Airfoil Data, NACA Report 824, 1945.

Loftin, L., K., Theoretical and Experimental Data for a Number of NACA 6A-Series Airfoil Sections, NACA Report 903, 1948.

Drela, M., XFOIL 6.9 User Guide, MIT Aero & Astro, 2001.

Green, Bradford, An Approach to the Constrained Design of Natural Laminar Flow Airfoils, NASA Contractor Report No. 201686, 1997.

Lutz, Th.,Airfoil Design and Optimization, Institute of Aerodynamics and Gas Dynamics, University of Stuttgart, 2000.

Garrison, P., The Shape of Wings to Come, Flying Magazine, November 1984.McGhee,R.,J., Viken, J.,K., Pfenninger, W., Beasley, W.,D., Harvey, W.,D., Experimental Results for a Flapped Natural-Laminar-Flow Airfoil with HighLift/Drag Ratio, NASA TM 85788, 1984.

Fujino, M., Yoshizaki, Y., Kawamura, Y., Natural-Laminar-Flow Airfoil Development for the Honda Jet, AIAA 2003-2530, 2003.

McCormick, B.,W.,Aerodynamics, Aeronautics and Flight Mechanics, 2nd Edition, John Wiley & Sons, New York, 1995.

Applied Aerodynamic Drag Reduction Short Course, University of Kansas Division of Continuing Education, Williamsburg, VA 1990.

Smith, A.,M.,O., High-Lift Aerodynamics, Journal of Aircraft, Volume 12, Number 6, June 1975.

Schetz, J.,A.,Boundary Layer Analysis, Prentice Hall, Upper Saddle River, New Jersey, 1993.

Cornish, J.,J., A Summary of the Present State of the Art in Low Speed Aerodynamics, Mississippi State University Aerophysics Department Internal

Memorandum, 1963.Raymer, D.,P.,Aircraft Design: A Conceptual Approach , AIAA Education Series, 1989.

Braslow, A.,L., Maddalon, D.,V., Bartlett, D.,W., Wagner, R.,D., Collier, F.,S., Applied Aspects of Laminar-Flow Technology, Appears in Viscous DragReduction in Boundary Layers, AIAA Progress in Astronautics and Aeronautics, Volume 123, 1990.

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