Laminar Flows 04
Transcript of 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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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-Laminar-
Flow Airfoil Development forthe Honda Jet.
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Natural Laminar Flow: Case Study
(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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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 = 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)
Fujino et al, Natural-Laminar-
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.