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
0.01
Cf
0.008
0.006
Laminar
0.004
Turbulent
0.002
0
1.0E+05
1.0E+06
Reynolds Number
1.0E+07
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 6series
NACA Report No. 903
Natural Laminar Flow
Comparison of NACA 6- and 6A-Series Pressure
Distributions
0.5
0
Cp
NACA 64-012
NACA 64-012A
-0.5
-1
0
0.2
0.4
0.6
0.8
1
x/c
XFOIL
• NACA 64-012: xtrupper = 0.5932, xtrlower = 0.5932
• NACA 64-012A: xtrupper = 0.6214, xtrlower = 0.6215
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.
2. Lutz, “Airfoil Design and Optimization,” 2000.
– Up to 65% laminar flow
3. Garrison, “Shape of Wings to Come,” Flying 1984.
– Drag as low as 30 counts
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-LaminarFlow 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 cruise
Fujino et al, “Natural-LaminarFlow 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-LaminarFlow 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-LaminarFlow 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-LaminarFlow 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 CLmax to be achieved
Ref: McCormick, “Aerodynamics, Aeronautics and Flight Mechanics,” pg. 202.
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
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
19541957
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
19631965
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
19851986
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
upper surface
lower surface
XFOIL Output
Why Does LFC Increases CLMAX?
• move boundary layer separation point aft
-1.0
(2196)
-0.25
(759)
-0.0625
(276)
-0.015625
(108)
0.0
m = 1/4
0.2
Cp
0.4
x0 = 1.0 ft
x0 = 0.25 ft
0.6
x0 = 0.0625 ft
0.8
Reynolds Number = 6x106
1.0
-1.0
-0.8
-0.6
-0.4
x0 = 0.015625 ft
-0.2
0.0
0.2
0.4
0.6
0.8
1.0
x - ft
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
-0.4
•Suction velocity required to maintain
incipient separation of the laminar b.l
and prevent flow reversal is given by:
adverse pressure gradient
-0.3
-0.2
vw  2.18  
-0.1
NACA 23012
cruise CL = 0.4
10,000 ft.
180 kts (303.6 ft/s)
Joseph Schetz, “Boundary Layer Analysis,” Equation (2-37)
0.0
0.0
0.5
leading edge
1.0
1.5
2.0
2.5
3.0
x (ft)
3.5
dU e
dx
4.0
trailing edge
45” x 12” grid – 439,470 holes
0.0025” dia
12” span
required suction velocity - vw (ft/s)
Suction Power Required for 23012 Cruise Condition
Preq = .00318 Hp / foot of span*
*assumes:
•use highest vw and Δp in calculation
•discharge coefficient of 0.5
•pump efficiency of 60%
0.035”
45” chord
Laminar Flow Control Approaches
1). Leading Edge Protection
required suction velocity - vw (ft/s)
2). Distributed Suction (perforated skin or slots)
-0.4
adverse pressure gradient
-0.3
-0.2
-0.1
NACA 23012
cruise CL = 0.4
10,000 ft.
180 kts (303.6 ft/s)
0.0
0.0
0.5
leading edge
Ref: Applied Aerodynamic Drag Reduction Short Course Notes,
…….Williamsburg,VA 1990.
1.0
1.5
2.0
x (ft)
2.5
3.0
3.5
4.0
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  7.5x106
•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 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
12.3m2/132.39ft2
2.877kN/647lbs
1360kg/2998lb
773 litres/194 USG
68kts (Clean) 55kts (Flap)
Drawing by
Anne Ocheltree & Nick Smalley
Wing & Tail
Calculating Laminar Flow
60%
Laminar
100%
Turbulent
Fuselage
40%
100%
Laminar
Turbulent
0.455
 0.0032
(log 10 Re) (1  0.144M 2 ) 0.65
Turbulent
Cf 
Lam C f 
1.328
 0.0005
Re
2.58
V-Tail
60% LM flow upper and lower surface
Fuselage Laminar to max thickness
Wing
60% LM flow upper and lower surface
2
Structure SWET (in )
Wing
224.89
Tail
58.39
Fuselage
295.87
SREF (in2)
132.72
Mcruise
0.29
Recruise 5.88E+06
Turb Cd
0.00875
0.00211
0.00975
Lam Cd
0.00268
0.00070
0.00473
% Reduction
69.41
67.05
51.51
Reduction in Drag from Laminar flow
0.025
0.02
Fuselage
Tail
Wing
0.015
Cd
0.01
0.005
0
Turb Cd
Lam Cd
Centuria NLF Manufacturing Tolerances
Rh,crit
h
hcrit (in.)
900
0.0072 inches
1800
0.0143 inches
2700
0.0215 inches
15,000
0.1195 inches
Carmichael’s waviness
criteria
0.0139 inch/inch

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, 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.