Laminar Flow

Rodney Bajnath, Beverly Beasley, Mike Cavanaugh

AOE 4124

March 29, 2004

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

CfLaminar

Turbulent

Natural Laminar Flow• NACA 6-Series Airfoils

– Developed by conformal transformations, 30 – 50% laminar flow

– Advantages: Low drag over small operating range, high Clmax

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

– Drag bucket: pressure distributions cause transition to move forward suddenly at end of low-drag Cl range

– Minimum pressure at transition location

NACA Report No. 824

Natural Laminar Flow

• NACA 6A-Series– 30 - 50% laminar flow– Eliminated TE cusp– Essentially same lift and

drag characteristics as 6-series

NACA Report No. 903

Comparison of NACA 6- and 6A-Series Pressure Distributions

-1

-0.5

0

0.5

0 0.2 0.4 0.6 0.8 1

x/c

CpNACA 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

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 on availability 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,” Flying 1984.

4. NASA Technical Memorandum 85788, 1984.

Natural Laminar Flow: Case Study

• SHM-1 Airfoil for the Honda Jet• Lightweight business jet, airfoil inversely designed,

tested in 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 for the Honda Jet.”

Natural Laminar Flow: Case Study (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 cruiseFujino et al, “Natural-Laminar-Flow Airfoil Development for the Honda Jet.”

Natural Laminar Flow: Case Study (Continued)

• Design Method– Eppler Airfoil Design and Analysis Code

• Conformal mapping, each section designed independently 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.”

Natural Laminar Flow: Case Study (Continued)

• Resulting SHM-1 airfoil– Favorable pressure gradient to 42%c upper

surface, 63%c lower surface– Concave pressure recovery (compromise

between Clmax, Cm, and MDD)

– LE such that at high α, transition near LE (roughness sensitivity)

– Short, shallow separation near TE for Cm

Fujino et al, “Natural-Laminar-Flow Airfoil Development for the Honda Jet.”

Natural Laminar Flow: Case Study (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 = 8x106 (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 = 10.3x106 (WT)

Fujino et al, “Natural-Laminar-Flow Airfoil Development for the Honda Jet.”

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 CLmax to be achieved

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

Notable Laminar Flow Control Flight Test ProgramsDate Aircraft Test Configuration LF Result Comments

1940 Douglas B-18

(NACA)

2-engine prop

bomber

NACA 35-215

10’x17’ wing glove section

suction slots first 45% chord

LF to 45% chord

(LF to min Cp)

RC = 30x106

Engine/prop noise effected LF

surface 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 span slots – both wings

full chord LF

RC = 47x106

effects of sweep on LF encountered

1985-1986

JetStar

(NASA)

4-engine business jet

two leading edge gloves

Lockheed – slot suction & liquid leading edge protection

McDD – perforated skin & and bug deflector

LF maintained to front spar through

two years of simulated airline

service

no special maintenance required lost LF in clouds &

during icing

LE protection effective

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

Why Does LFC Reduces Drag?• removes turbulent boundary layer

XFOIL output

Why Does LFC Reduce Drag?• turbulent boundary layer has a higher skin friction coefficient

XFOIL Output

upper surface

lower surface

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 = 6x106

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

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 in Low 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 in Low Speed Aerodynamics,” MSU Aerophysics Dept., 1963.)

Photographs Courtesy of the Raspet Flight Research Laboratory

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

required suction velocity - v

w (ft/s)

-0.4

-0.3

-0.2

-0.1

0.0

NACA 23012cruise CL = 0.4

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

adverse pressure gradient

dx

dUv e

w υ−−= 18.2

•Suction velocity required to maintain incipient 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

12

” sp

an45” x 12” grid – 439,470 holes

Preq = .00318 Hp / foot of span*

*assumes:•use highest vw and Δp in calculation•discharge coefficient of 0.5•pump efficiency of 60%

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

required suction velocity - v

w (ft/s)

-0.4

-0.3

-0.2

-0.1

0.0

NACA 23012cruise CL = 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.

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

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

Drag Benefit of Laminar Flow

CENTURIA• 4 Passenger Single Jet Engine GA Aircraft

• Competition•Cirrus SR22•Cessna 182

•Targets existing General Aviation pilots

•Cost ~ $750,000

•International Senior Design ProjectVirginia Tech and Loughborough University

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)

Drawing by Anne Ocheltree & Nick Smalley

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

==

=+

=

f

f

CLam

MCTurbulent

40% 100%

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 Cd % Reduction

Wing 224.89 0.00875 0.00268 69.41Tail 58.39 0.00211 0.00070 67.05

Fuselage 295.87 0.00975 0.00473 51.51

SREF (in2

) 132.72

Mcruise 0.29Recruise 5.88E+06

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

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

Carmichael’s waviness 0.0139 inch/inchcriteria

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

h

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

flow on wings, tail and fuselage

References•Abbott, 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 High Lift/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 Drag Reduction in Boundary Layers, AIAA Progress in Astronautics and Aeronautics, Volume 123, 1990.