# Aerofoil Design for Man Powered Aircraft

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1 Man Powered Aircraft Group Aerofoil Design for Man Powered Aircraft By F. X. Wortmann Universitat Stuttgart From the Second Man Powered Aircraft Group Symposium Man Powered Flight The Way Ahead 7 th February 1977 The Royal Aeronautical Society. London

2 From the Second Man Powered Aircraft Group Symposium, Man Powered Flight The Way Ahead. The Royal Aeronautical Society. London. 7 th February Aerofoil Design for Man Powered Aircraft By F. X. Wortmann Universitat Stuttgart 1. Design-objectives for MPA aerofoils Man powered aircraft are highly specialized machines not only from the structural point of view but also aerodynamically. They fly in a low Reynolds number range which is practically unexplored. On the other side, the requirements with respect to suitable aerofoils for the wing are much simpler than with sailplanes, since these machines are aimed to fly mostly between the best glide ratio and the stalling speed. For the usual high aspect ratio wings the aerofoils should develop a low drag minimum for C L values between 1.0 to 1.4 at Reynolds numbers from 3 x 10 5 to 7 x These numbers depend only slightly on the specific design. For the usual approach of splitting the total drag of the aircraft into a constant value C do and a lift dependent part the C L -value for the maximum glide ratio becomes C L2 opt = πc d0 + AR AR = Aspect ratio and for the minimum sinking speed C L2 opt = 3π C d0 AR In order to have a reasonable margin to the stall the C L max should be at least near 1.7 or 1.8. It is not so much the value of the C L max but the stall behavior which influences the flight safety. For this reason I would not recommend so-called maximum lift aerofoils /1/. Structural reasons ask for thick aerofoils. However, at these low Reynolds numbers aerofoil thickness has a much stronger influence on the drag than at higher Reynolds numbers, and their use should be restricted to the wing root. Since the necessary lift range is rather small the whole aerofoil problem can be reduced to the question: how can we attain low drag at these Reynolds numbers? I would like to simplify these matters even further and try to answer this question at first for symmetrical aerofoils. In a second step it is not too difficult to add a suitable camber and to look for an acceptable stall behavior. 2. Drag of symmetrical aerofoils at low Reynolds numbers Reliable drag measurements in the Reynolds number range of 3-7 x 10 5 are very rare. Therefore some exploratory tests were done in the laminar wind tunnel of the University of Stuttgart on some typical symmetrical aerofoils, which for two aerofoils go down to Re = 3 x In order to illustrate the all important transition control, two very different aerofoils have been selected and their velocity distribution at α = 0 is given in Fig.1. The first one, designated IS-30A/150, has a thickness of 15% and resembles in the front part a typical laminar aerofoil. Between 40 to 70% of the chord there exists a weak adverse pressure gradient called instability range /2/, which should destabilize the laminar boundary layer and provoke transition in front of the strong adverse pressure gradient. If this fails to work the laminar boundary layer will separate at the start of the strong pressure gradient at x/c = 0.68 The observed positions of the transition at several Reynolds numbers are indicated. It is somewhat surprising to see that the laminar boundary layer which

4 Fig.1 which alleviate the reattachment process and the start of the turbulent wall boundary layer. At the same time this type of velocity distribution is not sensitive to changes in Reynolds numbers and angles of attack and may be a practical solution. For the transition downstream of separation there seems not to exist any reliable empirical relationship. At least, three parameters are now involved: the momentum Reynolds number of the laminar boundary layer in the separation point, the level of preamplified perturbations at this point and finally the shape or thickness of the bubble, which seems to be correlated with the overall pressure gradient between separation and reattachment. At this point it may be appropriate to mention shortly the structural problems to obtain smooth aerofoil surfaces with thin membranes. The slight changes of the aerofoil contour due to the double curvature of the skin does practically not disturb the flow over straight wings, however, a wedge like nose or a kink in the chord wise contour is not tolerable. Therefore any type of construction should and could easily be tested with full scale models to ensure that laminar flow exists up to at least 70% of the chord on both sides of the model. The stethoscope is the cheap instrument to certify this fact. 4. Aerofoil design for the wing of man powered aircraft On the basis of the foregoing considerations four aerofoils with a thickness ratio of 12, 14, 16 and 18% have been designed. The shape of these aerofoils and the symmetrical aerofoil are given in Fig.6. Table 1 gives the coordinates. The thinner aerofoils are intended for the medium and outer wing sections. In this part of the wing the aileron effectiveness is important and a smaller thickness is in general favorable in this respect. At the same time the 12 and 14% aerofoils have additional characteristics which should improve the stall behavior. This feature is illustrated in Fig.7. At high angles of attack the velocity distribution on the upper side near the x/c =.7 station has a short region where the steep adverse pressure gradient is reduced to avoid a thick laminar separation bubble. and to favor the turbulent reattachment. The lower part of Fig.7 gives the curvature distribution, which also reflects this particular feature. It is the same type of boundary control as for low angles of attack, now applied at the nose to improve the high-lift case. It has been proven /5/ that this concept works even if the transition control cannot be perfect due to too low Reynolds numbers. With an improved initial development of the turbulent boundary layer on the upper side we can expect a separation position near the x/c =.7 position, which will not change very much with increasing angles of attack. Only the downstream angle of the separation region, which acts like a negatively deflected fluid flap, will increase, which in turn holds the pressure distribution over the forward part of the aerofoil and hence the overall lift nearly constant. It is hoped that at least some of these sections will be tested up to the time the lecture will be given. References /1/ F. X. Wortmann, The Quest for High Lift. AIAA Paper , MIT - Cambridge, Sept /2/ F. X. Wortmann, Experimentelle Untersuchungen an neuen Laminarprofilen fur Segelflugzeuge and Hubschrauber. Zeitschr. f. Flugwissenschaften 5(1957), S /3/ P.S.Granville, The calculation of viscous drag of bodies of revolution. David Taylor Model Basin Rep. 849(1953). /4/ E. R. v. Driest, Boundary layer transition. C. B. Blumer, AIAA J. 1963, Vol.I, pp /5/ F. X. Wortmann, Design of airfoils with high lift at low and medium subsonic Mach numbers. AGARD CP-102 (1972), Lisbon.

5 FX IS - 30A/ X/C 0.6 Fig. 1 Potential velocity distribution and transition position of two symetrical aerofoils at zero degrees angle of attack Re = Re = Re = Re = Re =

6 Cd IS-30A/150 FX FX e Fig. 2 Drag values of three symmetrical aerofoils with 15, 12 and 10% thickness ratio at alpha = zero degrees for Reynolds numbers between 300,000 and 2,000,000 Re cd 5

7 CL Re 6 10 Re Re Cd alpha X t/c Transition Fig. 3 Drag polars, CL alpha and transition position of the FX at low Reynolds numbers

8 CL FX IS-30A/ Re Fig. 4 Drag polars of the IS-30A/150 and the FX at Reynolds numbers near one million Cd

9 Re v Transition Instability Separation Flat Plate H 32 Fig. 5 Momentum thickness Reynolds number of instability and transition as function of the shape parameter H32 1 Granville 2 V. Driest 3 Delata Rev = 800

10 FX76MP-160 FX76MP-140 FX76MP-120 FX Fig. 6 Shape of some aerofoils designed for low Reynolds numbers

11 V/V infinity FX 76-MP120 Alpha = 8.7 degrees CL = X/C SQRT (X/C) 0.4 C/R X/C SQRT (X/C) 0.4 Fig. 7 Velocity and curvature distrabutions of the FX 76-MP120 near the nose

12 2.0 Institut fur Aero - und Gasdynamik der Universitat STUTTGART Laminarwindkanal FX76MP-120 Re = 700,000 Re = 500,000 C Re = 300,000 C L M CD alpha Xtr/c

13 2.0 Institut fur Aero - und Gasdynamik der Universitat STUTTGART Laminarwindkanal FX76MP-140 Re = 700,000 Re = 500,000 C Re = 300,000 C L M CD alpha Xtr/c

14 2.0 Institut fur Aero - und Gasdynamik der Universitat STUTTGART Laminarwindkanal FX76MP-160 Re = 700,000 Re = 500,000 C Re = 300,000 C L M CD alpha Xtr/c

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