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 NATIONAL ADVISORY COMMITTEE 
 FOR AERONAUTICS 
 
 TECHNICAL MEMORANDUM 1331 
 
 INVESTIGATIONS OF THE BOUNDARY-LAYER CONTROL ON A FULL 
 
 SCALE SWEPT WING WITH AIR BLED OFF 
 
 FROM THE TURBOJET 
 
 By P. Rebuffet and Ph. Poisson-Quinton 
 
 Translation of "Recherches sur l'Hypersustentation d'une Aile en 
 
 Fleche Reelle par Controle de la Couche Limite Utilisant le 
 
 Prelevement d'Air sur le Turbo -Reacteur." la Recherche 
 
 Aeronautique, O.N.E.R.A., No. 14, March-April, 1950. 
 
 Washington 
 April 1952 
 
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 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
 
 TECHNICAL MEMORANDUM 1331 
 
 INVESTIGATIONS OF THE BOUNDARY -LAYER CONTROL ON A FULL 
 
 SCALE SWEPT WING WITH AIR BLED OFF 
 
 FROM THE TURBOJET* 
 
 By P. Rebuffet and Ph. Poisson-Quinton 
 
 SUMMARY 
 
 The following account reviews the various stages of a research 
 program relative to the high-lift devices on a swept wing by combined 
 suction and blowing (jet action), with ejectors fed by air bled off 
 (extracted) from the turbojet. 
 
 After reviewing the essential principles of the boundary-layer 
 control obtained by comparison with theory, the electric analogies and 
 the wind-tunnel tests as well as the essential elements of ejector 
 operations, the writers describe the tests made in the large tunnel at 
 Chalais-Meudon on a full-scale model of the SO 6020 wing. 
 
 They comment on some results relative to take-off and landing and 
 examine the airplane -turbojet energy balance. 
 
 The writers emphasize that this investigation has been carried out 
 in a remarkable spirit of teamwork and particularly their chief collab- 
 orators Messrs. Mirande and Ravailhe, Jousserandot and Chevallier. 
 
 They have also drawn on the unpublished reports of Messrs. Lebrun, 
 Legras, Nieviaski, Ponteziere of the O.N.E.R.A. and of Mr. Chiaffiote 
 
 of the Hispano-Suiza Company. 
 
 NOTATION 
 
 S reference area 
 
 Vq speed at infinity 
 
 *"Recherches sur l'Hypersustentation d'une Aile en Fleche Reelle par 
 Controle de la Couche Limite Utilisant le Prelevement d'Air sur le 
 Turbo-Reacteur." la Recherche Aeronautique, O.N.E.R.A., No. 1^, March- 
 April, 1950, pp. 39-5^. 
 
NACA TM 1331 
 
 s width of blowing slot 
 
 Z local chord 
 
 I geometric incidence 
 
 i aerodynamic incidence 
 
 q v , q m , q flow volume, flow mass, and flow weight 
 
 V Q mean speed of suction flow tube coefficient of flow 
 
 q 
 
 c q = 7Zr~ (expanded to 15° C, 760 mm of mercury) 
 SV 
 
 C q 'j C "j C flow coefficient of injection suction and blowing 
 
 C^ momentum flow coefficient 
 
 c 
 
 area of mixer cross section _ ^m 
 area of injector cross section S' 
 
 area of diffuser outlet section S 
 
 q m v £ 
 
 2 S 
 
 a 
 
 area of mixer section Sm 
 
 L __ length of mixer (circular ejector) 
 D diameter of mixer 
 
 u _ mass flow of suction fluid _ ^ m 
 mass flow of engine fluid q 1 
 
 momentum flow at blowing slot ^m V 
 momentum flow at ejector " q' m V' 
 
 e = 
 
 flow of kinetic energy at blowing slot ^m Y^ 
 flow of kinetic energy of engine jet q' m y t 2 
 
 INTRODUCTION 
 
 The airfoils suitable at sonic speeds are generally characterized 
 by a low C Zjmx , which is still further reduced by the sweptback form 
 of the wing. 
 
NACA TM 1331 
 
 These two factors led to the study of high-lift devices designed 
 to develop high C z on modern high-speed aircraft, to improve the 
 landing as well as the take-off. 
 
 The control of the boundary layer responsible for the separation 
 of flow and the limitation of Cz has been studied for a long time on 
 thick airfoils, without combining the suction or the jet action with a 
 modification of the airfoil camber. The power involved and the flows 
 (suction or blowing) were considerable. 
 
 Numerous technical studies have been made especially by the 
 Germans (ref . 1) during the war, on more suitable airfoils and with 
 partial-span flaps; some were flight -tested. 
 
 In France, certain airplane firms included in their research 
 program on high-lift devices, the study of suction or jet action, and 
 it was proposed to combine the two methods by utilizing injectors. 
 The S.N.C.A.S.O. had considered, since 19^6, the construction of a 
 scale model of the large tunnel at Chalais-Meudon, for the SO 6008 bis 
 airfoil. 
 
 The jet airplane affords, moreover, a new possibility of boundary- 
 layer control by suction or by jet action (blowing). 
 
 The first solution has been the subject of several studies 
 (refs. 2, 3)> the second, more recently, was tested by the Hispano- 
 Suiza firm. 
 
 The aerodynamical department of the O.N.E.R.A. has, since 19^-6, 
 correlated the research data on boundary -layer control, checked the 
 wind-tunnel data against those obtained in potential flow by electric 
 analogy, and made a theoretical study of the effect of sinks (suction). 
 
 These two comparisons made it possible to orientate the experimental 
 investigations, since the purpose of the boundary -layer control is to 
 assure a potential flow under certain provisions to be specified later 
 on. 
 
 These basic researches had to be made on a Chalais-Meudon mock-up 
 which permitted: 
 
 Operation at a Reynolds number near that at landing, the mock-up 
 having the dimensions of a flying airplane 
 
 Testing of components at full scale 
 
 Allowance for certain number of contingencies imposed by a product 
 similar to that of a real airplane 
 
NACA TM 1331 
 
 It was advisable to make this study on a concrete case. The 
 earlier projects of the S.N.C.A.S.O. were computible with the program 
 of the O.N.E.R.A., the control devices were applied to a swept wing 
 model of the SO 6020 type. Besides, bleeding off air from the turbojet 
 made it possible to reproduce at Chalais-Meudon, the image of the air- 
 plane, by dissociating first the airplane from the turbojet, which was 
 installed outside of the tunnel, thus, assuming a safer operation as 
 well as a more accurate way of recording the drag, that was to be 
 determined simultaneously with the lift. 
 
 It was, nevertheless, possible to measure the thrust of the jet, 
 with due regard to the air bleeding from the turbojet. 
 
 The study at Chalais-Meudon had to be made in competition with the 
 O.N.E.R.A. through: 
 
 The S.T.A., engine section which had put the Nene Hispano-Suiza 
 turbojet at the disposal of the O.N.E.R.A. 
 
 The Arsenal de 1' Aeronautique, which had lent the turbojet support 
 frame and the corresponding control cabin 
 
 The Hispano-Suiza Company that constructed the collector for the 
 air bleed and assumed the control of the turbojet for the dura- 
 tion of the tests 
 
 The S.N.C.A.S.O. which, manufactured the model at the request of 
 the O.N.E.R.A. and made the study and perfected it 
 
 Lastly, the airplane section of the S.T.A. which underwrote the 
 tests Chalais-Meudon 
 
 About 45 to 50 men were involved in this undertaking, which 
 included: 
 
 51^ polars and pitching moments 
 
 2000 hinge moments 
 
 31500 pressure measurements on the wing (700 pressure gage 
 
 negatives) 
 85O photographs 
 speed and temperature measurements 
 
NACA TM 1331 
 
 PRINCIPLES AND APPLICATION OF BOUNDARY- LAYER CONTROL 
 BY HIGH-LIFT DEVICES 
 
 Of the multitude of systematic studies by the O.N.E.R.A. on this 
 subject (ref . h) simple guiding ideas can be enumerated: 
 
 It is known that the maximum lift of an airfoil is limited by the 
 break away of the boundary layer from the upper surface. This separa- 
 tion has its origin in the appearance of a pressure gradient and is 
 accentuated with the incidence or with the deflection of a flap. The 
 remedy is to reduce the peaks of the speed increase by acting on the 
 curvature of the center line of the profile or else to prevent or limit 
 this separation by controlling the boundary layer. 
 
 In the first method, the fluid is assumed perfect and yields an 
 incurved profile center line, in a kind of downward deflection of 
 several flaps, hinged front and rear. 
 
 The Peres -Malavard electric analogy method affords a very fruitful 
 study of the best distribution of the increase of speed over the upper 
 surface of an airfoil) for a given lift coefficient, it is possible to 
 increase the camber of the mean line for distributing the peaks of the 
 speed increase accompanying each variation incurvative due to the flap 
 setting. 
 
 The most dangerous increase, which appears at the leading edge can 
 be reduced by a deflection of the latter which puts the stagnation point 
 back at the nose of the airfoil. This deflection must be such that the 
 increase at the nose and the droop are of the same order of magnitude. 
 
 In viscous fluid the maximum lift is limited by the separation 
 downstream from the peaks) it can be remedied by boundary-layer control 
 either by suction or blowing. Both methods have a secondary effect on 
 the circulation around the airfoil: source effect for suction which is 
 theoretically calculable, or the induction effect for blowing. It has 
 been shown that the lift increase was proportional, respectively, to 
 the induced flow and to the momentum flow. This secondary effect, 
 negligible as long as the fluid is separated, remains apparent only 
 after elimination of the separation. In other words, plotting the 
 lift increase ££7. against the flow parameter Cq or the momentum 
 
 Cjj, yields a curve with asymptotic direction from the critical flow 
 assuming the readherence of the boundary layer. Above this value, the 
 profile can be regarded as functioning in perfect fluid for which the 
 laws {£Cz, flow) are linear. Experience indicates that it is not 
 important to exceed the readherence flow from the energy point of view, 
 because the lift increases only very little. 
 
NACA TM 1331 
 
 The first wind-tunnel test was made on wings with drooped nose 
 fitted with flaps and slotted suction or blowing slots in the droop. 
 Experience has proved the efficiency of the boundary-layer control for 
 obtaining readherence, the maximum lift being no longer limited except 
 by the separation at the leading edge. 
 
 The applications are difficult owing to the over-all dimensions 
 of the suction or blowing channels and to the practical impossibility 
 of mounting auxiliary compressors in the airplane. To make it practi- 
 cable, it is necessary to utilize a part of the engine power installed 
 on the airplane and to reduce the dimensions of the air ducts along 
 the span. 
 
 This is the reason why our studies on the application of boundary- 
 layer control were oriented from the beginning toward the employment 
 of compressed air, making it possible to distribute the blast along 
 the wing span through a small size duct supplied by air drawn from the 
 compressor of the turbojet. 
 
 The compressed air can be expanded directly in a blowing slot but 
 this method poses problems difficult to solve: discontinuity of stressed 
 skin covering at the upper surface, control of the width of a very fine 
 slot before being rigid, and especially the correct distribution of the 
 flow along the wing span. 
 
 Lastly, the ejector, in spite of its low energy efficiency, has the 
 advantage of simplifying the wing structure considerably and at the same 
 time assuming the boundary-layer control by suction followed by blowing 
 at two points of the profile. 
 
 EFECTORS 
 Definitions 
 
 In general, an ejector (fig. 1) consists of: 
 
 An injector supplying the "primary air" characterized by the 
 velocity head p 1 (dynamic pressure) 
 
 A suction chamber and suction cone where the sucked or "secondary" 
 air enters 
 
 A mixer whose speed and temperature are equalized 
 
 A diffuser terminating at the blowing slot 
 
WACA TM 1331 
 
 It is characterized by the geometric parameters A. and <x; its 
 aerodynamic functioning is defined by u, g, and e. 
 
 Ejectors have been studied in static tests as well as on airfoils 
 in wind tunnels for a long time. It is conceded that the static pres- 
 sure existing on the profile at the location of the suction and jet 
 action are similar, the functioning of ejectors is the same with or 
 without flow past the profile, this remark justifies the static tests. 
 
 The jet velocities have, for the utilization of ejectors, as far 
 as the boundary-layer control device is concerned, as much significance 
 as the suction volume) in such ejectors the operating conditions differ 
 considerably from the classical. 
 
 Tests 
 
 The theoretical study and the tests involved several types of 
 ejectors: an ejector with flat mixer, without diffuser and a set of 
 ejectors with circular mixer terminating in a diffuser, the section of 
 which was developed from a circle to a rectangle (fig. l) . 
 
 The parameter X can be varied by changing the section of the 
 mixer or by varying the section and the number of injectors; the 
 parameter a by modifying the mixer or the diffuser. The effect of 
 the mixer length L and that of the variation of the loss of lead due 
 to suction was studied at the same time. 
 
 The comparison of the two types of ejectors was extended to the 
 coefficient of discharge u, the speed of the blowing jet Vs, and the 
 ratio of the momentum g. The results prove the identity of the 
 functioning of the flat and circular ejectors. 
 
 The choice between the two solutions can be guided by considera- 
 tions of size, weight, and facility of construction. 
 
 The circular ejectors favor high values of X; values that give 
 
 high discharge u but a low momentum ratio g, with blowing slots of 
 
 little thickness, because the section of the mixture does not occupy 
 the entire span of the wing. 
 
 The flat ejectors with flat injectors simplify the design, but 
 produce only small values of X for these straight slots, hence a 
 low u, but a high g. 
 
 The best solution therefore seems to lie in a combination of 
 ejectors with flat mixers and multiple circular injectors, so as to 
 obtain high X. values for straight blowing slots. 
 
NACA TM 1331 
 
 Arrangements Adopted at Chalais-Meudon 
 
 The arrangements and the dimensions of the ejectors for the model 
 of the SO 6020 airplane used in the large wind tunnel at Chalais-Meudon 
 was determined in the light of these tests. It was finally decided to 
 adopt multipler circular ejectors with the following characteristics: 
 
 X = 28 a = 0.75 |=5 
 
 Figure 2 defines the development of the curves u, g, e as 
 functions of the injection pressure p 1 . A brief study of this graph 
 shows that the induced flow is about 2l times that of the injected flow, 
 
 that the momentum flow at the injector exit is sensibly maintained, but 
 that the ratio of the kinetic energies is relatively small, of the order 
 of 20 percent. 
 
 In the range of operations customary at Chalais-Meudon, the injec- 
 tion pressure p' is such that the speed in the throat of the injectors 
 is sonic . 
 
 The characteristic coefficients u, g, and e depend, to some 
 extent, on the loss of lead at suction and, consequently, on the setting. 
 
 At g = 0.8 for the straight slot corresponding to a^ = 0, 
 g reaches in effect, unity for the a-j_ = 15 setting. 
 
 It is quite evident that it is important to design the suction slot 
 correctly in order to obtain correspondingly minimum loss of lead. 
 
 Exploration of Blowing Jet on a Flap of the Wing SO 6020 
 
 In this study the injectors were supplied by compressed air for 
 the purpose of ascertaining the velocity distribution in the blowing jet. 
 
 In the case of the circular ejectors, figure 3 shows at 100 mm 
 downstream from the blowing slot a regular distribution in span already, 
 although with a slight fluctuation attached to each ejector. 
 
 This distribution is still more uniform when using an ejector with 
 flat mixer and circular injectors, of a total section equal to that of 
 the preceding arrangement but having twice the number of injectors. 
 
 Figure k shows the velocity profile of the blowing jet in the axis 
 of an ejector or injector 100 mm downstream from the blowing slot. 
 
NACA TM 1331 
 
 TESTS ON A METAL MODEL IN TWO-DIMENSIONAL FLOW 
 
 Before making the tests on the full-scale model in the large tunnel 
 at Chalais-Meudon, a very complete study was made on a SO 6003 bis, 
 metal airfoil of 0.838-m chord, 1-m span, equipped with ejectors 
 (fig. 5)- According to a static test for defining the ejector charac- 
 teristics, followed by a test between panels in the wind tunnel at 
 Porte d'Issy, this wing element showed very promising results from the 
 point of view of high-lift devices for landing and take-off. 
 
 The pressure distributions measured on the profile have been 
 compared with those obtained in the electric tank (fig. 6). 
 
 The identity of the curves assumes that the separations are com- 
 pletely reabsorbed by blowing, while this is not the case for high 
 settings. Moreover, suction and blowing introduce a secondary effect 
 of sinks and induction that was not represented in the theoretical 
 study. It follows a different chordwise load distribution for an 
 identical lift. 
 
 Figure 7 refers to the following configuration: 
 
 leading edge setting r\ = 30° 
 
 first suction flap a-^ = 25 
 
 second blowing flap clq = V?° 
 
 which are marked 30 B-25-^5° > the letter B indicates that, the leading- 
 edge slot is plugged . 
 
 At constant incidence the readherence of the boundary layer to 
 the first and second flap is manifested by a sudden increase in C z ; 
 beyond the flow of readherence the linear law (C z , C^) prevails, which 
 corresponds to the action on the potential flow. 
 
 The unit curves (C z , i) undergo a translation of the ordinates with 
 increasing Cu, but it seems that the incidence of C Zmax decreases 
 
 as a result of the stalling of the leading edge which proceeds from a 
 critical value of the maximum speed increase connected with the leading- 
 edge radius. 
 
 A series of tests was run on this wing element with varying leading- 
 edge radius. 
 
10 NACA TM 1331 
 
 For a given contour, the stall produced for a given C z is so 
 increased considerably as the leading-edge radius is increased. The 
 analysis of the pressure distributions shows the great increase of 
 admissible critical speed increases, when the radius increases. 
 
 The solution of the combined flat mixer and multiple circular 
 injectors was recently attempted on the metal wing element SO 6008 bis. 
 
 The model was tested in the wind tunnel at Cannes. According to 
 the first results obtained, there is a net improvement of C^ (of the 
 order cf 30 percent), which is manifested by a more effective control 
 of the boundary layer. 
 
 The practical construction of flat mixers reveals itself, as 
 predicted, much easier than that of multiple circular ejectors. 
 
 THE SO 6020 MODEL AT CHALAIS-MEUDON 
 
 The 10-percent thick wings copy the outside form of those of the 
 SO 6020; they are joined to a fuselage, modified only in the front and 
 rear parts, the length of the actual fuselage being unsuitable for the 
 wind tunnel. 
 
 The construction is of wood, but with due regard to the placement 
 of the essential components of the real wing structure. 
 
 The wing has an adjustable leading edge and two flaps (fig. 10), 
 the setting a,j_ is controlled by electric jets and measured by 
 teledynes. 
 
 Cemented strain gages afforded the over-all measurement of the 
 hinge moment of the flaps a^_ + 09. 
 
 A series of static pressure orifices was installed in the section 
 AB (fig. 8). Exploratory graphing hooks of the boundary layer are 
 arranged at both sides of slot a^ and in the blowing jet a^. The 
 wool streamers on one wing enabled the flow to be checked, and in 
 particular to follow the readherence on the different parts. 
 
 Compressed Air for Model 
 
 The turbojet is housed with its control cabin outside of the wind 
 tunnel (fig. 11). The air is bled off through special pipes which 
 deflect a part of the flow feeding the combustion chambers (fig. 12) 
 and terminate in a double collector. 
 
NACA TM 1331 11 
 
 The air enters the model through a symmetrical circuit (fig. 8) 
 without introducing momentum, with insertions of flexible elements, 
 that cause no stress at the wind-tunnel balance. 
 
 At entry in the model, a special joint enables variations of the 
 incidences without vitiating the measurement of the pitching moment. 
 
 The characteristics p 1 , T' of the fluid (sensibly at rest) 
 before being distributed to the injectors spaced along the span, and 
 also the total flow injected, q'p^ are measured. 
 
 In all tests the temperature T' is about 50° C, whereas, during 
 bleeding off it may reach 200° C. For equal mass flow, the speed of 
 ejection and hence the momentum flow is therefore much similar. The 
 experimental results must be corrected for application to the airplane. 
 
 Functioning of the Turbojet 
 
 Figure 13, made at the Hispano-Suiza test stand, shows the evolu- 
 tion of the principal characteristics of the turbojet plotted against 
 the pressure and the flow bleed-off weight. 
 
 The tests at Chalais-Meudon were carried out by varying the speed 
 of the turbojet between 6000 and 11,600 rpm. The experimental points 
 corresponding to three speeds are shown on the graph, they are situated 
 on a straight line defined by the total section of the injectors. 
 
 The diagram contains the network of constant rotational velocities 
 and bleed-off temperatures, the isotemperatures in the nozzle (after 
 the turbine) and, in particular, that which limits the diagram to its 
 upper part. The isothrust curves of the turbojet also included, involve 
 the heat balance. 
 
 Procedure of Test 
 
 For a given geometrical form of the model, at an incidence I and 
 a tunnel speed V~o, the inlet pressure at the injectors (and consequently 
 the primary flow of the ejectors) were varied by changing the speed of 
 the turbojet. The airspeed Vq was varied between 18 m/sec and 35 m/sec 
 
 (3 X 10 6 < Re < 6 X 10 6 ). 
 
 To each injected flow q'-nj there corresponds a secondary flow 
 uq' which is none other than the flow sucked through the slot a^j 
 
 the total flow (l + M-)q' p is ejected by the blowing slot a^. The 
 momentum flow is decreased and from it the parameter C^. . 
 
12 NACA TM 1331 
 
 The balance measures the three components of the resultant} the 
 elements necessary for the calculation of the pitching moments, hinge 
 moments, and pressure distributions are indicated; the observation 
 and the wool tuft photographs are extremely instructive. 
 
 Reference Area 
 
 For computing the lift and drag coefficients, the wing area 
 affected by the boundary-layer control was chosen, as it affords a 
 much easier comparison with the tests in two-dimensional flow. 
 
 RESULTS OF TESTS 
 Representative Parameters of the Aerodynamic Phenomena 
 
 Suction and blowing induce flow changes at the wing which are 
 manifested by lift and drag changes. 
 
 Lift . - Boundary-layer control makes it possible to make the flaps 
 more efficient, and hence to increase the lift considerably. Moreover, 
 the suction entails an increase in circulation resulting in a £C Z = K. 
 Z£q", the coefficient K being considerably higher as the suction slot 
 
 at the upper surface is nearer to the trailing edge. In our specific 
 case (slot at 57-5 percent from leading edge), this coefficient, given 
 by the theory, is very low (K = 2.38) and the gain in lift is practically 
 
 negligible for the flows involved (C n " « — — ) . By contrast the rise in 
 
 \ q 100/ 
 
 the circulation by the induction of the blowing jet entails a con- 
 siderable increase in the lift Z£z = K' /C^j the coefficient K 1 like- 
 wise increases when the jet approaches the trailing edge, but it cannot 
 be defined theoretically; the effect of blowing being preponderate, it 
 is reasonable to represent the evolution of C z as function of the jet 
 action parameter C^ . 
 
 Drag . - The suction-blowing combination must be considered: A 
 certain mass of fluid q" m is sucked in at a mean speed Va (it can 
 be obtained by measuring the velocity distributions above the lip (or 
 rim) upstream from the slot); a mass q m is ejected through the blowing 
 slot at a speed V s . 
 
NACA TM 1331 13 
 
 At low incidences for which the speeds V s and V a can be con- 
 sidered identical with their projections on the speed at infinity Vg, 
 we obtain a force 
 
 -Rx = 
 
 <3m v s " q" m Va = L 1 + ^ )V s " V a]« 
 
 m 
 
 v a 
 
 (1 + H) 
 
 ^m 
 
 in the absence of an exact measurement of V a , it can be estimated equal 
 to Vq with due allowance for the local increase of speed at this 
 
 position (fig. lk) , on the other hand, the V s ~ UVq and \x = 3- The 
 subtractive term is thus very small and is neglected. 
 
 Hence 
 
 q V or 2 
 _ ^m s _ ^q _ 
 
 x " 1 2 " s/z " H 
 | P SV ^ s ' L 
 
 s/z being the relative width of the blowing slot. The drag curves are 
 likewise plotted against C^, which has been computed from Cq 1 by 
 
 C 
 
 2C q ,2 (l + u) 2 
 
 with due amount of the variation of \i with the pressure. 
 
 Velocity distribution near the suction and blowing slots . - Fig- 
 ure Ik shows by way of example a velocity distribution upstream and 
 downstream from the suction slot as well as downstream from the blowing 
 slot, for several values of C^; the boundary layer upstream from the 
 first slot is little affected by the suction, and completely suppressed 
 downstream from it, the pocket of the speeds of the blowing jet indicates 
 on the other hand, that the induction effect accelerates the fluid well 
 above the jet itself. 
 
ik NACA TM 1331 
 
 Mechanism of Readherence of Boundary Layer 
 
 It is interesting to follow the successive stages of readherence 
 of the flow over the rear part of a wing for a specific configuration 
 of flaps: nose not drooped, first slot suction aj_ = 25°> second slot 
 blowing, a-2 = ^4-5°, incidence I = 12°. 
 
 The pressure distributions recorded in the section AB are shown 
 in figure 15, the visualization of the flow by wool tufts in figure l6, 
 and the lift curve (I = 12°) as function of C^ in figure 17. For 
 C|_l = 0, the two flaps have completely separated (phase a). The first 
 flap readheres rather quickly when suction begins, while the flow is 
 still severely agitated over the second flap (100 Cp. = 1.3> phase b); 
 nevertheless, the increment of the lift is appreciable. 
 
 In the next phase (c), corresponding to 100 C^ = 5»5> the flow 
 
 has completely separated from the two flaps. The orientation of wool 
 tufts indicates the complete suppression of the oblique flow toward 
 the tip of the sweptback wing. These different phases are manifested 
 in the pressure distribution by a marked increase of speed over the 
 rear part of the profile, which tend toward the theoretical increases 
 of speed, whereas the correlative increase of circulation entails an 
 increase in the speed at the leading edge. If the flow is increased 
 further, there comes a moment where the circulation is such that the 
 maximum increase at the leading edge exceeds the critical value and 
 induces separation. In this phase (d), the lift increases is no longer 
 linear and the curve (C z , Cp) bends inward. When the separation from 
 the wing is generalized (I = 15° and over), another lift increase due 
 to the action on the potential flow and the component of the blowing 
 momentum for increasing flows, is observed. 
 
 The formation of the total drag as function of C^ is also • 
 represented in figure 17; it increases with C^ as a result of the 
 increased induced drag. 
 
 If the latter is curtailed, the curves, for which the flow is 
 sound, are sensibly reduced to a unique curve the mean slope of which 
 is that of the theory (-C x = +Cu). 
 
 Figure 18 shows that the readherence for the C^ is considerably 
 
 greater as the profile curvature is more pronounced; nevertheless, all 
 curves tend toward one asymptotic direction. 
 
 A drooped nose reduces the speed increase at the nose as seen from 
 the pressure distributions of figure 19- Its effect on C Zmax is 
 discussed later. 
 
NACA TM 1331 15 
 
 Cruising Configuration 
 
 The static tests proved that the aerodynamic resultant is equal to 
 the blowing momentum; this roughly checks with the result obtained in 
 the study of ejectors. 
 
 The same property is again verified with relative wind as seen in 
 figure 20, where the setting of the polar s is sensibly equal to C^ at 
 incidences for which the flow, without boundary-layer control, remains 
 sound. 
 
 With boundary-layer control, the polars are sensibly parallel to 
 the induced polar, although the latter is computed by the classical 
 method of Prandtl, not applicable rigorously, to a swept wing. 
 
 Take-Off Configuration 
 
 Different configurations with moderate flap settings are designed 
 for use at take-off. Certain forms are to be discarded for a specific 
 take-off speed, be it that they do not furnish the corresponding C z , 
 or give an incidence too close to separation or give rise to abnormally 
 high drag. A range of take-off C 2 is established for seeking the 
 best forms chosen by the minimum (C x - C x j_) . Figure 21 shows the 
 corresponding network; it is apparent that high Cz necessitate a more 
 pronounced setting. 
 
 Figure 22 compares the drag of different forms at constant C z 
 and variable o 2 . The C x = f(ag) at the right for C^ = and 
 100 Cp. = 5 is shown by way of example. 
 
 For 02=0, the gain in C x is greater than C^, and remains 
 substantially the same when 02 varies. 
 
 Landing 
 
 The influence of the three elements that define the profile camber 
 is decomposed, in order to obtain the best Cz max . 
 
 In the study of the effect of the drooped nose, the continuity of 
 the profile was reestablished by an appropriate setting. Figure 23 
 shows the role of the drooped nose which, by reducing the speed increase 
 at the nose, delays the separation and consequently lengthens the unit 
 curves C z = f( i) . 
 
16 NACA TM 1331 
 
 The best setting t) lies between 15° and 30 . A study was made 
 also on a drooped nose the suction being assured through the wing by 
 ejectors in similar manner to the suction on flap a]_, whose slot was 
 then closed (fig. 10). 
 
 It could be observed at this occasion that the flap ai overcame 
 the separation, even for a setting of 25 , by induction of the blowing 
 jet. 
 
 The suction in the droop of the nose, by impeding the separation 
 the forwf 
 
 a little too. 
 
 from the forward part of the wing, postpones the incidence of C z 
 
 Varying only a-|_, the best setting increases with Cq, owing to 
 the prevented separation due to the suction, according to the curves 
 of figure 2k. 
 
 The effect of the blowing flap cl^ (fig- 25) shows that the 
 optimum is near k0° or ^5° > depending on the values of C^ . 
 
 dC„ 
 In addition, — - increases considerably with C^; this factor is 
 d<X2 
 
 favorable for assuming the lateral control by blowing aileron. 
 
 For a contour of primary interest from the point' of view of landing 
 it is advisable to take the total drag that corresponds to Cz mov for 
 
 several C^ into consideration. 
 
 Figure 26 gives by way of example the polars obtained for the con- 
 figuration 30 B-25-^5°- The rise in C z is substantial, especially 
 for a swept wing with thin profile. Correlatively, the very significant 
 drags are due in a large measure, to the induced drag. 
 
 Pitching Moments and Hinge Moments 
 
 Pitching moments . - One difficulty accompanying the use of high- 
 lift devices is the correlative appearance of a high pitching moment. 
 The equilibrium of the airplane necessitates an appreciable elevator 
 setting which becomes strikingly nonlifting and reduces the Cz max of 
 
 the airplane to the same extent. 
 
 The tests made at Chalais-Meudon on a model without empennage do 
 not permit a direct treatment of the problem. 
 
NACA TM 1331 17 
 
 The O.N.E.R.A. has recently begun a series of studies in the wind 
 tunnel at Cannes, on a complete airplane model equipped with drooped 
 nose and blowing flap. The first results indicated that polars 
 balanced up to C z of the order of k are obtainable without having to 
 change the initial setting of the tail. 
 
 Incidentally, it is noted that the blowing effects are favorable 
 on the tail that is in a sound flow and highly deflected; the only 
 limitation is the stalling over the tail. 
 
 Hinge moments . - The flaps a^ and ag are constructed in four 
 elements distributed along the span of a semispan wing (fig. 8). The 
 hinge-moment measurement of a-|_ and a^ results in the following: 
 
 For the contour 30 B-25-^5° taken as example, the hinge moment 
 decreases 30 percent from root to tip, when the control of the boundary- 
 is inactive. 
 
 The improvement of the flow over the suction and blowing flaps, 
 multiplies the initial hinge moment by a factor varying from 2 to 3-5, 
 from root to tip. This confirms the qualitative conclusions drawn 
 from the wool tuft observations, particularly in the case of swept wing. 
 
 The hinge moment does not change much with the incidence, even a 
 little beyond C Zmax , an element favorable for lateral control. 
 
 For a nondrooped nose, the hinge moments start to decrease before 
 C Zmax , even with boundary-layer control; this confirms the significance 
 of adopting a drooped nose. 
 
 As to the forces required for flap control, they do not seem to 
 pose particular structural problems. 
 
 BALANCE AIRPLANE TURBOJET 
 General 
 
 Bleeding off air from the turbojet limits its thrust as a result 
 of the increased temperature upstream from the turbine. It must be 
 allowed for in the performance determinations. 
 
 By contrast, the decrease in thrust for a given turbojet speed 
 is negligible as shown in figure 13, while the specific consumption 
 increases a little during the short period of bleed-off. 
 
18 NACA TM 1331 
 
 The preliminary study of ejectors and the over-all results pointed 
 out previously have brought into evidence a relationship between the 
 injected and blown momentums . 
 
 Consequently, with the employed notations 
 
 oV = ss m v s = gP^ |sv 2 
 
 The aerodynamic study defining C^, the weight and the C z defining 
 
 Vq, the term q'^V supplied by the turbojet, is deduced. It is 
 
 important to locate the corresponding points of operation on the bleeding- 
 off diagram. 
 
 To this end, a net of curves of equal q ' V * can be plotted on the 
 chart from which the equivalent surface of the injectors is then deduced. 
 
 Landing 
 
 By way of example, for a specified landing speed of h-5 m/sec and 
 operation at 100 C^ = 10, we get q' m V' = 266} on the extrapolated 
 
 diagram of the turbojet, the maximum thrust limited by the temperature, 
 is I78O kg equal to C x = 0.67) if the airplane C x is higher, the 
 permanent flight is slightly lower; for the usual wing loading, the rate 
 of descent is not more than 1 to 2 m/sec . The diagram also shows that 
 the injector exit should be considerably increased. 
 
 Take -Off 
 
 The adaptation to an optimum configuration is obtained by a similar 
 process: for a specified Cz = 1-75, the configuration is OB-15-15 
 (fig. 21)) the maximum climb is reached with 100 C^ = 2.5, the corre- 
 sponding drag is balanced by a thrust P = 1028 kg) the excess thrust 
 (AP = 2190 - 1028) makes it possible to obtain the rate of climb. 
 
 The diagram shows the ejection section to be much smaller than that 
 used on the Chalais-Meudon model. 
 
 CONCLUSIONS 
 
 The foregoing results demonstrate the effectiveness of boundary- 
 layer control on a wing of pronounced sweep and whose basic profile is 
 comparatively thin. 
 
NACA TM 1331 19 
 
 They show, moreover, that the source of energy used, bleeding off 
 
 from a turbojet, results in a significant energy balance since the 
 
 thrust loss is negligible at take-off and enables landings in sensibly 
 horizontal steady state. 
 
 The summary numerical applications merely prove that different 
 numbers of injectors must be used at take-off and landing, such as two 
 rows of injectors, for example, one for take-off, both for landing. 
 
 A study under way on suction at shock wave level, on airfoil 
 SO 6008 bis, zero incidence, has proved the complete absence of break- 
 away for extremely small suction flows . 
 
 It seems justifiable to elaborate the study of the solution by 
 ejectors and flat mixers. A much greater control efficiency and simple 
 design should be attainable, since the mixer is reduced to two flat 
 surfaces . 
 
 The lateral control should be examined correlatively with the 
 high-lift devices. The efficiency of a differential aileron setting 
 should be checked, an efficiency which appears attained, since precisely 
 the flow over the flap cig is correct, even at high incidence. 
 
 Lateral control by differential injection might be explored. 
 
 Such studies have been scheduled in the supplementary program on 
 the Chalais-Meudon model. 
 
 Although the work at Chalais-Meudon was made on a full-scale wing, 
 with corresponding limitations, it should be pointed out that only 
 laboratory studies are involved, and that the application of boundary- 
 layer control by bleeding off from the turbojet poses technical and 
 structural problems which are outside the scope of the present research 
 program. 
 
 Translated by J. Vanier 
 National Advisory Committee 
 for Aeronautics 
 
20 NACA TM 1331 
 
 REFERENCES 
 
 1. Poisson-Quinton: Idees nouvelles sur le controle de la couche limite 
 
 applique aux ailes d' avion. Les Cahiers d'Aerodynamique, no. 3> 
 19^5- 
 
 Deplante: Rypersustentation, commandes Transversales . Technique et 
 Science Aeronautique, no. 2, 19^+6. 
 
 2. Sedille: La propulsion par reaction en combinaison avec 1' aspiration 
 
 de la couche limite. Ier Congres National de 1' Aviation, 19^5* 
 
 3- Dupin et Morain: L' alimentation en air des turbo-re'acteurs par 
 
 aspiration de la couche limite. Technique et Science Aeronautiques, 
 no. 5, 19 Vf- 
 
 k. Poisson-Quinton: Recherches theoriques et experimentales sur le 
 controle de la couche limite. Vile Congres International de 
 Mecanique Appliquee, Londres, septembre 19^3. 
 
WACA TM 1331 
 
 21 
 
 Injector 
 
 Suction 
 cone 
 
 5l ~S 
 
 Diffuser 
 
 Surfoce 
 
 discharge area S 
 
 Figure 1.- Test setup of ejector. 
 
 0.175 
 0.15 
 
 0.5 
 
 X -~ 28 
 
 a I 0.75 Suction s,ot a7% ^ 15 
 
 1000 
 
 1500 
 
 2000 p .mm Hg 
 
 Figure 2.- Parameter of operation of the ejector units used on the S.O. 6020 
 
 wing. 
 
22 
 
 NACA TM 1331 
 
 /Ejector 8^ 
 
 Circular ejectors (ejectors 6tos] 
 
 X = 28Cf=0J5~ 
 • Flat mixer (injectors 1 1 to 16) "^^ 
 
 X= 44 a = 0-75 
 100 2O0mm 
 
 Vm 
 
 /s ' 
 
 Figure 3.- Spanwise distribution of blowing jet. 
 
NACA TM 1331 
 
 23 
 
 f : Slot width of flap a 2 
 I -.Circular nozzle t,= 5 mm 
 2: Flat mixer f 2 = 9 mm 
 
 Figure 4.- Velocity profile 100 mm downstream from blowing slot. 
 
2k 
 
 NACA TM 1331 
 
 Figure 5.- Metal model; view of inside, cover of flap a^ has been removed. 
 
NACA TM 1331 
 
 25 
 
 ♦ c, 
 
 8 
 
 -7 
 
 0B-I5-45 
 
 \z± 
 
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 L 
 
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 * 
 
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 * 
 
 Figure 6.- Comparison of theoretical and experimental pressure. 
 
26 
 
 NACA TM 1331 
 
 -^ „!poc^-- 5 
 
 I00C^.= 2 
 
 iboc^,= 
 
 Figure 7.- Example of high-lift device in two-dimensional flow. 
 
NACA TM 1331 
 
 27 
 
 Sweepback y?=3i° 20' 
 1/4 chord line 
 
 Figure 8.- Model S.O. 6020. 
 
28 
 
 NACA TM 1331 
 
 Figure 9.- Setup in large tunnel at Chalais-Meudon. 
 
NACA TM 1331 
 
 29 
 
 Figure 10.- Section of profile: suction from first flap or the leading edge. 
 
30 
 
 NACA TM 1331 
 
 Figure 11.- Turbojet and control cabin. 
 
 Figure 12.- Bleeding off air from turbojet. 
 
NACA TM 1331 
 
 31 
 
 Figure 13.- Characteristics of Hispano-Suiza Nene turbojet operation with 
 
 bleed -off. 
 
32 
 
 NACA TM 1331 
 
 Figure 14.- Velocity profile adjacent to suction and blowing slots. 
 
NACA TM 1331 
 
 33 
 
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3^ 
 
 NACA TM 1331 
 
 Figure 16.- Visualization: (a) flaps separated, (b)ai partially separated, 
 
 (c)a., and cig readhering. 
 
NACA TM 1331 
 
 35 
 
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36 
 
 NACA TM 1331 
 
 Figure 18.- Effect of flap setting on the C M of readherence. 
 
NACA TM 1331 
 
 37 
 
 Figure 19.- Effect of drooped nose. 
 
38 
 
 NACA TM 1331 
 
 Figure 20.- Flap polar not set. 
 
NACA TM 1331 
 
 39 
 
 Figure 21.- Comparison of take-off configurations. 
 
ko 
 
 NACA TM 1331 
 
 4052 
 
 Figure 22.- Effect of blowing flap setting on drag. 
 
NACA TM 1331 
 
 kl 
 
 Figure 23.- Effect of drooped nose t). 
 
k2 
 
 NACA TM 1331 
 
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NACA TM 1331 
 
 ^3 
 
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 1000^=13 
 
 Figure 26.- Family of polars for a landing configuration. 
 
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 UNIVERSITY OF FLORIDA 
 DOCUMENTS DEPARTMENT 
 120 MARSTON SCIENCE LIBRARY 
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