Laminar Flow

Rodney Bajnath, Beverly Beasley, Mike Cavanaugh

AOE 4124 March 29, 2004

Introduction

• Why laminar flow?

– Less skin friction Lower drag 0.01

0.008

0.006

0.004

0.002

0 1.0E+05

Skin Friction: Laminar vs. Turbulent

1.0E+06

Reynolds Number

1.0E+07 Laminar Turbulent

Natural Laminar Flow

• NACA 6-Series Airfoils – Developed by conformal transformations, 30 – 50% laminar flow – Advantages: Low drag over small operating range, high C lmax – 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 C l 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

Natural Laminar Flow

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

0.5

0 NACA 64-012 NACA 64-012A -0.5

-1 0 0.2

0.4

0.6

0.8

1

x/c

• NACA 64-012: x tr upper = 0.5932, x tr lower • NACA 64-012A: x tr upper = 0.6214, x tr = 0.5932

lower = 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 1. NASA Contractor Report No. 201686, 1997.

– Up to 65% laminar flow 2. Lutz, “Airfoil Design and Optimization,” 2000.

– Drag as low as 30 counts 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 C lmax – 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

– C lmax = 1.6 for Re = 4.8x10

– Loss of C l 6 , M = 0.134

less than 7% due to contamination – C m > -0.04 at C l = 0.38, Re = 7.93x10

6 , M = 0.7

– Airfoil thickness = 15% – M DD > 0.70 at C l = 0.38

– Low drag at cruise Fujino 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 C l max 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 C lmax , C m , and M DD ) – LE such that at high α, transition near LE (roughness sensitivity) – Short, shallow separation near TE for C m Fujino et al, “Natural-Laminar Flow Airfoil Development for the Honda Jet.”

Natural Laminar Flow: Case Study (Continued)

• Specifications: – C lmax = 1.66 for Re = 4.8x10

– 5.6% loss in C lmax 6 , M = 0.134

due to LE contamination (WT) – C m = -0.03 at C l = 0.2, Re = 16.7x10

6 (Flight) – C m = -0.025 at C l = 0.4, Re = 8x10 6 (TWT) – M DD – M DD – C d – C d = 0.718 at C = 0.707 at C = 0.0051 at C = 0.0049 at C l l l l = 0.30 (TWT) = 0.40 (TWT) = 0.26, Re = 13.2x10

6 = 0.35, Re = 10.3x10

6 (TWT) (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 Boundary layer thins and becomes fuller across slot outer skin plenum chamber inner skin 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 C Lmax to be achieved Ref: McCormick, “

Aerodynamics, Aeronautics and Flight Mechanics

,” pg. 202.

Date

1940 1955 1954 1957 1963 1965 1985 1986 Notable Laminar Flow Control Flight Test Programs

Aircraft Test Configuration LF Result Comments

Douglas B-18 (NACA) 2-engine prop bomber Vampire (RAE) single engine jet F-94 (Northrup/USAF) jet fighter X-21 (Northrup/USAF) jet bomber 30° sweep JetStar (NASA) 4-engine business jet NACA 35-215 10’x17’ wing glove section suction slots first 45% chord upper surface wing glove suction - porous surface full chord suction NACA 63-213 upper surface wing glove suction – 12, 69, 81 slots new LF wings for program suction through nearly full span slots – both wings LF to 45% chord (LF to min C p ) R C R C = 30x10 6 full chord LF M~0.7 / R C =30x10 Full chord LF 0.6 < M < 0.7

R C = 36x10 6 full chord LF = 47x10 6 6 Engine/prop noise at M effected LF surface quality issues Monel/Nylon cloth 0.007” perforations local >1.09 shocks caused loss of LF effects of sweep on LF encountered 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 upper surface lower surface XFOIL Output

Why Does LFC Increases C

LMAX

?

• move boundary layer separation point aft

-0.0625

(276) -0.015625

(108) -1.0

(2196)

0.0

-0.25

(759) m = 1/4

0.2

C p

0.4

0.6

x 0 = 1.0 ft x 0 = 0.25 ft x 0 = 0.0625 ft

0.8

1.0

-1.0

Reynolds Number = 6x10 6 -0.8

-0.6

-0.4

-0.2

x 0 = 0.015625 ft

0.0

0.2

x - ft

0.4

0.6

0.8

1.0

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 •C LMAX = 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 •C LMAX = 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 -0.4

-0.3

adverse pressure gradient

•Suction velocity required to maintain incipient separation of the laminar b.l and prevent flow reversal is given by: -0.2

-0.1

NACA 23012 cruise C L = 0.4

10,000 ft.

180 kts (303.6 ft/s) 0.0

0.0

0.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

leading edge x (ft) trailing edge

v w

  2 .

18  

dU e dx

Joseph Schetz, “Boundary Layer Analysis,” Equation (2-37) 0.0025” dia 45” x 12” grid – 439,470 holes P req = .00318 Hp / foot of span* *assumes: •use highest v w and Δp in calculation •discharge coefficient of 0.5

•pump efficiency of 60% 45” chord 0.035”

Laminar Flow Control Approaches 1). Leading Edge Protection Ref: Applied Aerodynamic Drag Reduction Short Course Notes, …….Williamsburg,VA 1990.

2). Distributed Suction (perforated skin or slots) -0.4

adverse pressure gradient

-0.3

-0.2

-0.1

NACA 23012 cruise C L = 0.4

10,000 ft.

180 kts (303.6 ft/s) 0.0

0.0

0.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

leading edge x (ft) trailing edge 3). Hybrid Laminar Flow Control

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  7.5x10

6 •Cruise C L  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 Project Virginia Tech and Loughborough University

Centuria Design Details

• Cruise altitude • Cruise Speed • Range • Take-off run • Aspect Ratio • Wing Area • Thrust • MTOW • Fuel Volume • Stall Speed 10,000ft 185kts 770nm 1575ft 9.0

12.3m

2 /132.39ft

2 2.877kN/647lbs 1360kg/2998lb 773 litres/194 USG 68kts (Clean) 55kts (Flap)

Drawing by Anne Ocheltree & Nick Smalley

Wing & Tail Calculating Laminar Flow Laminar 60% Turbulent 100% Fuselage Laminar 40% Turbulent 100%

Turbulent Lam C f

 (log 10 0 .

455 Re) 2 .

58 ( 1  0 .

144

M

2 ) 0 .

65  0 .

0032

C f

 1 .

328 Re  0 .

0005

Fuselage Laminar to max thickness V-Tail 60% LM flow upper and lower surface Wing 60% LM flow upper and lower surface

Structure

Wing Tail Fuselage

S REF (in 2 ) M cruise Re cruise S WET (in 2 )

224.89

58.39

295.87

132.72

0.29

5.88E+06

Turb C d

0.00875

0.00211

0.00975

Lam C d

0.00268

0.00070

0.00473

% Reduction

69.41

67.05

51.51

Reduction in Drag from Laminar flow

0.025

0.02

Cd

0.015

0.01

0.005

0

Turb Cd Lam Cd

Fuselage Tail Wing

Centuria NLF Manufacturing Tolerances

R h,crit h crit (in.) 900 0.0072 inches 1800 0.0143 inches 2700 15,000 0.0215 inches 0.1195 inches  h Carmichael’s waviness 0.0139 inch/inch criteria Ref: A.L. Braslow, “Applied Aspects of Laminar-Flow Technology,” AIAA 1990

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

, 2 nd 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.