MMlA-\w°\ JUi brary Battelle Memorial Institute Columbus, Ohio MED JW • - NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS TECHNICAL MEMORANDUM No. 1209 EXPERIMENTAL STUDY OF FLOW PAST TURBINE BLADES By E. Eckert and K. v. Vietinghoff-Scheel Translation of "Versuche uber die Strbmung durch Turbinenschaufelgitter" Vorabdrucke aus Jahrbuch 1942 der deutschen Luftfahrtforschung, 6. Lieferung Washington UNIVERSITY OF FLORIDA Tnno 1 cue DOCUMENTS DEPARTMENT June 1949 1 2Q MARST0N SC | ENC E LIBRARY P.O. BOX 117011 GAINESVILLE, FL 32611-7011 US/> NATIONAL ADVISORY COMMITTEE FOE AERONAUTICS TECHNICAL MEMORANDUM NO. 1209 * EXPERIMENTAL STUDY OF FLOW PAST TURBINE BLADES By E. Eckert and K. v. Vietinghoff-Scheel OUTLINE I. INTRODUCTION II. ESTIMATION OF PRESSURE DISTRIBUTION OVER TURBINE BLADES III. THE INTERFERENCE METHOD IV. EXPERIMENTAL LAYOUT V. EXPERIMENTAL RESULTS VI. SUMMARY AND OUTLOOK I. INTRODUCTION The requirements on gas turbines for aircraft power units, namely, adequate efficiency, operation at high gas temperatures, low weight, and small dimensions, must he taken into consideration during the design of the blading. To secure good efficiency, it is necessary that the gas flow past the hlades as smoothly as possible without separation. This is relatively easily obtainable in the accelerated flow of turbine blading, if the blade spacing is cho3en small enough. A small blade spacing, however, is detrimental to the other requirements outlined above. Operation at high gas temperatures usually calls for blade cooling. This cooling is associated with a power input that lowers the turbine effi- ciency. Since the amount of heat that must be carried off for cooling a blade can be influenced rather little, the gross power input Tor a tur- bine 3tage can be reduced by keeping the number of blades to a minimum, that is, with blades of high spacing ratio. But here also a limit is imposed, the exceeding of which is followed by separation of flow. Hence the requirement of finding blade forms on which the flow separates at rather high spacing ratios. Small dimensions of the turbine are essentially obtained by keeping the outside diameter of the blading as small as possible. This is made possible by choosing a high tip speed and making the width of the annular space of the turbine stage available for the passage of the ga3 great, that is, the inside diameter as small as possible. But on such long blades the flow at the inside diameter is appreciably different from that at the outside diameter. The flow strikes the rotor blades inside at a *"Versuche uber die Stromung durch Turbinenschaufelgitter. " Vorabdrucke aus Jahrbuch 19^2 der deutschen Luftfalirtforschung, 6. Lieferung, pp. 2—10. NACA TM No. 1209 much flatter angle than outside. The turbine blades must therefore ensure a flow free from separation throughout the whole available range of flow angles. Since the visualization of the separation phenomena on a running turbine involves considerable difficulties, it is appropriate to study the flow first on blade grids at rest. Direct application of such findings to the guide vanes of a turbine is probably possible. Greater care is advised in application to rotor blades. On these the boundary layer, which form3 at the surface of the blades, is under the influence of the centrifugal force. This should not affect the separation phenomena on axial turbines very much, where the centrifugal force is perpendicular to the direction of motion. A greater effect is to be expected on radial turbines, where the centrifugal force acts against the separation when the gas flows through the turbine from the inside toward the outside. In the reversed flow direction the centrifugal- force effects favor separation of flow. The use of the interference method for the study of flow past the blade grids (references 1 and 2) has the advantage that the tests can be run at Eeynolds and Mach numbers encountered on actual turbines. On top of that, interference photographs not only afford a qualitative picture of the flow process, but also can be interpreted quantitatively, such as the determination of the pressure distribution over the blades and with it the torque exerted on the rotor, for instance. II. ESTIMATION OF PRESSURE DISTRIBUTION OVER TURBINE BLADES To the torque exerted by the flowing gas on the rotor there corre- sponds a force on each blade in circumferential direction. This force is introduced by low pressure at the back of the blade and by high pres- sure at the face of the blade, in the same manner as the lift on an air- foil. Weinig (reference 3) computed the pressure distribution caused "by the flow of a frictionless fluid for several blade forms, one of which is represented in figure 1, by the pressure distribution curve for a blade grid with spacing ratio — = O.96 and flow direction w-, . The viscosity of the flowing gas causes a boundary layer along the blade surface, and it is this, as is known, that effects the separation at the blade as 3oon as the pressure rise in flow direction along the blade surface exceeds a certain value. To be feared most of all is a sepa- ration at the back in the area between A and B, although it can occur equally n the face at C. The amount of pressure rise which a boundary layer can overcome without separation depends upon whether the layer is laminar or turbulent. It Is therefore to be expected that the separation phenomena discussed hereinafter are affected by the Reynolds number. The order of magnitude of the pressure rise to be overcome by the flow past a turbine blade is readily estimated when the entrance and exit NACA TM No. 1209 angle of flow for the blading are known. Weinig's calculations (fig. l) as well as the test data on wings and cascades of airfoils (references k and 5) give the pressure distribution curve an approximately triangular form. Giving the pressure distribution projected on a straight line (width of grid b in fig. 2) perpendicular to the grid direction the shape of a triangle as represented in figure 3, the magnitude of the force U per unit blade length exerted by the flow in grid direction is V,^f (1) where Ap is the maximum pressure difference between front and back of blade, and b the width of the blade grid. On the other hand, by the momentum theorem with the notation of figure 2A this force U follows the equation U = paw m (w u2 - w-ui) (2) the flow velocities being assumed so small that constant gas density p can be a33umed. Since an estimate is involved, this is possible up to the speed of about two thirds the velocity of sound. This assumption simplifies the calculation, which, of course, can be carried out for variable density just as well. The velocity components w are intro- duced vectorially in equation (2), that is, subtracted when in the same direction, but added when directed oppositely. Determining the maximum pressure difference from equations (l) and (2) gives Ap m = 2p|w m (w u2 - w ul ) (3) The stage pressure drop Ap , is the difference of the static pressures p-j_ and p 2 upstream and downstream from the blade grid as Ap 3t = Pi - p 2 = |(w 2 2 - w x 2 ) = |(w u2 2 - w ui 2 ) (k) The dynamic pressure of the flow velocity is 11 = fa 2 (5) With the 3e quantities the pressure distribution triangle can be plotted to scale, as exemplified in figure 3A, because the pressure at the blade **• NACA TM No. 1209 trailing edge, that is, at the tip of the triangle, is practically p ? . The amount of the pressure rise at the "back of the "blade follows then as the difference between po and the lowest pressure p, as P 2 " P 3 = Ap m " 4 1 " Ap st < 6 > Whether the "boundary layer is able to overcome this pressure rise without separation depends upon the magnitude of the kinetic flow energy relative to this pressure rise. The velocity increases from w-, to w 2 during the flow through the blade grid. For the present estimate, therefore, it is best and perhaps accurate enough as well to refer the pressure rise p p - p.-, to the average kinetic energy of flow (7) With equation (6) we then get Po - P- a w m(%2 " w ul) -x 2 2 2 w u2 " w ul ■ 2 2 2 w-, + W2 b 2 2 w-, + w 2 2 2 w l + w 2 (8) as measure for the danger of separation. The pressure distribution triangle in figire 3A relates to a turbine 3tage with axial entry (w u ^ = 0) and an exit angle p ? of 20° ( w m/ w u2 = ^ an Pp = 0.36M niea-sured against the circumferential direction. Eeversing the flow through the blade grid involves a pressure rise from p toward p-j_. The blades must then, of course, be shaped a little differently, the pointed tip of the blade being placed on the exit side again, as exemplified in figure 2B. Equations (3), (4), and (7) remain valid for this direction of flow. But the dynamic pressure of the flow velocity is 12 = |w 2 2 (9) KACA TM Wo. 1209 The pressure distribution triangle assumes the form represented in figure 3B, and the referred pressure rise at the blade back amounts to 2 2 w u2 " w ul , ,* Z3 = 8 a-q*u2-"ul)_ ^2 +2 ^ _ b 2 2 ? 2 2? w 2 + w l w l + w 2 w l + w 2 The flow directions of figure 3B ar © the same as in figure 3A- A compari- son of these two triangles indicates that the pressure rise, referred to the average kinetic energy of flow which the boundary layer has to over- come, is substantially higher for the blower blade than for the turbine blade. This is the reason such marked deflection as assumed in figure 3 can apparently nob be achieved at all in a blower blading without sepa- ration. According to equations (8) and (10 ) the danger of separation is so much greater as the flow is more deflected and the length ratio a/b is greater. The result is that a turbine stage can be operated with a greater ratio a/b, hence with a greater spacing rati^ a/t or with a greater deflection of flow, than a blower stage. The present estimate gives obviously only an approximate picture of the separation danger since the effect of blade form^ and profile chord, for example, are not mentioned at all. But even so it serves as a guide for the evaluation of grids of highly cambered blades, for which there are practically no test data available. The preparation of more accurate data on profile form is the purpose of the present report. The maximum lift coefficient c & max is not suit- able for evaluating the separation danger on highly curved blades in cascade arrangemsnt, because existing constant pressure turbines operate o at lift values up to 6, while the flow over an airplane wing already separates at c a max «w 1-5 • On the airplane wing there exists a definite relat' onship between Vo - p^ lift coefficient c a and the parameter ^ inti educed here. On ^Reviewer's note: This equation does not agree with Figure 3B. P2 - P3 Pi " ?3 In order to do so, should be changed to -. i q That closely spaced, highly curved blades, for which the flow is very considerably dependent upon the blade form, can result in a pres- sure distribution other than triangular is shown in figure 9° 2 a / w u2 - w uA u The relation for blade grids is c„ = - 2 j where Wo, Is the value of the vectorial mean of the two velocities w^ and w 2 - 6 NACA TM No. 1209 substituting a triangle for the pressure distribution plotted against the profile chord t, the lift A per unit of span is A = Ap m i, while, by definition, A = c a tq. Lastly: * p 2 ' p 3 p 2 - P 3 = Ap m - q.; hence i = 2c a - 1. P 2 - P. The parameter ~ 2 is therefore equivalent to the maximum 1 lift coefficient of wings c «* 1.5 • III. THE INTERFERENCE METHOD The tests were made with a Mach-Zehnder interferometer (reference 6) manufactured by Zeiss, (fig. k) . It consists essentially of four plane mirrors, two of them (&]_ and agl being "half- silvered," and the other two (b-, and ^q) with opaque silver coating. Every beam of light from the light source c, is split into two parts at the half-silvered plate a-,, and reaches the screen e by different paths. One part is reflected at a-,, reaches mirror b 2 , passes through mirror a 2 and reaches screen e with a portion of its intensity. The beam going through mirror a, arrives after reflection at the mirrors b-j_ and a 2 at the screen e. The two light rays passing through the interferometer when superimposed produce interference fringes at the point of intersection (reference 1) . If the four mirrors are perfectly parallel the two light rays leaving plate &„ intersect at infinity. Thus the interference fringes would occur at infinity. Since the wave fronts of the two light rays are parallel to each other, however, the width of the interference bands is infinitely great. Bands of finite width at infinity are obtained when the two mirrors a-, and a are turned through a small angle a-, and a-p out of their neutral position. The plane in which the interference bands originate can be shifted to any position at infinity and the width itself can be adjusted as desired by corresponding choice of angle a~ and a^. At the setting of the mirrors shown in figure k, the two light rays move in divergent directions from plate a^. Their extension backward meets in the plane 1-2, however, so in this plane a virtual interference pattern is produced. This picture can be made visible on screen e by a converging lens d and photographed. In reality, a more complicated optical device takes the place of the lens. For the subsequent application of the interferometer the plane 1-2 must be placed in the position between mirrors a„ and a-, ■^Reviewer's note: This equation is correct only if q is changed to q^. If that is done the lift coefficient c & is based on q 2 which is unusual but not necessarily incorrect. NACA TM No. 1209 shown in figure k. If the axea of rotation of ' mirrors a, and a~ are normal to the plane of the drawing, the interference fringes thrown on the screen are parallel to the axes of rotation. By swinging mirror a^ about a second axis which lies in the mirror plane and is normal to the first axis, the direction of the bands can he varied at will. Their relative spacing is varied by the adjustment of the angles a-, and ccg. Monochromatic light produces contrasting interference fringes in a larger field. For the present purpose a mercury vapor lamp with a monochromatic filter which lets through light of the wavelength \ m = 0. 56^+1 X 10"3 milli- meter was employed. Placing a chamber f closed by two flat parallel windows g and filled with a gas in the path of the rays and varying the density p of the gas in the chamber by a value p ' , the interference pattern shows a shift of the bands. The density variation in the chamber Ap = p ' - p can be computed from the observed band shift by the formula (reference l) Ap = °— (11) L(n - 1) X is the wavelength of the light in vacuum, e the band shift measured in widths (one width equals the distance of the center lines of two suc- cessive light and dark bands), n the index of refraction of the gas of density, p and L the path- length of the light rays in the medium of n - 1 density p . The expression has a constant value for every gas. P n 1 J+ For air Z = 0.002265 — — -. The path length L for the setup used P kg s 2 is given by the inside distance between the two windows g. It amounts to 199« 8 millimeters. These values entered in equation (ll) give 2 Ap = 0.001246 e \ S (12) m Given the type of change of state by which the density variation p' - p of the gas in the chamber f was obtained, all the other conditions of state can be computed from the fringe displacement. For the isentropic change of state of an ideal gas it is p'/p = (p '/p) K and T'/T = (p'/p)*" , for example. Through the density p' computed by equation (ll) the 8 NACA TM No. 1209 pressure p 1 and the temperature T" are defined. Likewise, for an isobaric change of state the gas equation £ = ET gives the temperature T" = 11-2- P In place of chamber f the test section of a flow channel closed at both sides by parallel windows g is placed in the path of the rays. The blades to be studied are fitted into this channel in such a way that the light rays pass parallel to the generating axis of the blades. Through the air flow a density field is formed around the blades. Now the density has a different value for the path of each light ray through the channel. The result is a distortion of the interference fringes. The blade grid, itself, together with the interference fringes originating in the plane 1 - 2, is reflected on the screen e. Therefore, the density field around the blades can be determined by measuring the deflections of the interference fringes at each point of the screen. As is seen from the subsequent photographs the density field of the flowing air made thus visible brings out the extent of the boundary layer as well as its separation. IV. EXPERIMENTAL LAYOUT The flow channel into which the blade grid was mounted is shown in two sections in figure 5> ancL in photograph figure 6. The air is induced by a blower through the rectangular inlet cone a and flows past the blading b. The air jet leaving the screen is intercepted by the exit cone c and returned to the blower by way of the diffuser d and a pipe line connected by a flexible leather collar. The blower is driven by a direct- current motor so that its speed can be controlled within wide limits. It produces a maximum pressure difference of 260 millimeters of water. Since the flow through the blades was to be explored at different flow angles, the front - and back wall f of the air channel before the blading are pivotable about the two rotational axes g. The blade grid could be investigated for 3ix flow directions. Every setting called for a different entrance cone a. The two parallel windows i were mounted in the heavy side walls h of the channel. The front - and back wall f was sealed from the sides h by rubber collars and plasticine. The blades b are suspended from two tension wires k of 2 millimeter gage and sealed from the glass windows by glued on rubber washers Z. To prevent the exit direction of the air flow from the blading from being influenced by the position of the exit cone c, the dead air regions m and n existing at either side of the jet were joined by two strong pipe lines, by which a pressure balance is maintained between the dead air regions. In spite of this the uniformity of the air jet behind the blade grid is so far still not quite satisfactory, and is to be improved on the newly designed set-up by larger dead air regions m and n. The blade3 WACA TM No. 1209 were impregnated "beech wood. Great care was taken to ensure the beat possible two- dimensional blade form, (flat, parallel generating line), because the light rays must pass parallel to the blade aa exactly as possible over its entire length if an observation of the flow processes in the thin boundary layer of the blade surface is to be possible. Appro- priate gages ensure exact parallel setting of the blades during assembly on the tension wires. The grid spacing was varied by changing the number of blades. The surfaces were finished with shellac to ensure smoothness. In figure 7 two blades together with the tension wires are reproduced. In the majority of the tests described hereinafter the middle blade of the screen was heated. To this end the blade was provided with two holes o into which a chromium-nickel heating coil on a ceramic tube was inserted. Owing to the low thermal conductivity of the wood it was, of course, not to be expected that the blade surface would reach a constant temperature, which, however, did not matter in the present tests. The exact alinement in the light rays of the interferometer was obtained by means of three set screws p. From the 1^-6 millimeter diameter circle presented by the window for inspection, the interferometer covers a rectangular field of view of 8 x 10 centimeters. The position of the field can be changed by shifting the test section. The flow velocity w Q of the blades is computed from the negative pressure Ap measured at the orifice r relative to the test room, by the equation v o 2 Ap = \_2_ (13) 2g A. = density of air. The upper wall f of the channel contains further three orifices s closed by threaded plugs, into which a small pitot tube can be inserted, for checking the uniformity of air flow in front of the blades. Up to a thin boundary layer at the side walls the velocity over the cross section was practically constant and agreed with the figure obtained by equation (13) to a few tenths of one percent. A check of the flow behind the blades by pitot tube is afforded by a flange, shown in figure 5> replacing one of the two glass windows. The flange consists of a ring t and a concentric disk u which can be turned by means of a handle v. Its setting is read from a scale w. This disk carries a hole x for inserting a pitot tube. By moving the pitot tube in this orifice and turning the disk u the flow in the section behind the blades can be measured. This is important, in order to ascertain whether the flow separates from the side walls h. This phenomenon wa3 repeatedly observed on blower blade grids (references 5 an "two to four blades were mounted in the working section. This made spacing ratios a/t, (fig. 2), of O.687, O.859, and 1.141 possible. The chord of the blade profile is 58.2 millimeters. It was measured, as for a wing, as projection of the blade profile on a straight line touching the lower surface of the blade, figure 1. Through the different settings of the channel walls f and the related entrance cones flow angles of 20, 3^, U8, 62, 76, and 90° were obtained. Figure 8 shows an interference photograph at O.687 spacing and 3^ flow angle. The deflection of the interference fringes caused by the density field in the flowing air is plainly visible. Directly at the blade surface the interference lines have a bend, that is, an especially great density gradient exists at the surface-^, which is due to the fact that the boundary layer of the flow is heated a3 a result of the heat (heat of dissipation) liberated by internal friction. This phenomenon makes it possible, as already indicated by Th. Zobel, to render the bound- ary layer visible by interference photographs. Even the dead air region behind the blade trailing edge, which is formed by the warmer air in the boundary layer, is clearly shown. Outside of the boundary layer and the dead-air region the air flow is, of course, free from loss, the change of state of the air in this area is therefore isen tropic, hence the pressure field within the flow can be computed from the density field defined by the interference photographs by means of the previously cited relations. This method also yields the pressure distribution along the blade surface from the fringe displacements at the border of the boundary layer, as exemplified by the pressure area, figure 9, where the pressure distribution is plotted against the grid width b (fig. 2), hence against the blade projection on a normal to the grid direction. The area of the pressure surface indicates the tangential force acting on the blade. The photograph, figure 8, was made at 21.7 meters per second air speed. The mean exit velocity at the end of the blade channel computed from the continuity equation was 89.5 meters per second. The mean velocity at the same point from the interference record is 91-6 meters per second. The blower operated at its maximum speed. The fringe movements (fig. 8) are not yet very great at these velocities, and the accuracy with which the tangential force on the blade can be determined from the interference photographs is, as a result, not quite satisfactory. This drawback is easily removed by addition of a stronger blower. The boundary layers at the blade inlet side where the velocities are comparatively low, are not clearly visible in figure 8. ^Unfortunately, many details on the photographs are not as clear as on the originals. MAC A TM No. 1209 11 Since for the study in question, however, it is important to observe the separation of the boundary layer, even at the low velocities, the temperature of the boundary layer was raised artificially by heating the blade. According to the Navier-Stokes differential equations for frictional fluid flow the field of flow is not affected by temperature differences if the fluid properties (density, viscosity) are not dependent upon the temperature and the lifting forces in the flow introduced by the temperature differences disappear with respect to the inertia forces. And this is certainly the case at the speeds and low increases of temperature of 25 at the mo3t produced within the boundary layer by the heating^. Figures 10 to 30 represent interference photographs with such blade heating. The two heavy parallel lines are the shadows of the current supply wires, carried in thin insulating tubing along the upright side walls of the entrance cone, figure 6. The flow direction is nearly the same as the direction of these wires. Figures 10 to 15 show the flow at the smallest spacing ratio a/t = O.687, figures l6 to 21 at a/t = O.859, and figures 22 to 27 at the greatest spacing ratio a/t = l.l4l. Eeverting to figures 10 to 15 it is seen that for this smallest flow angle the flow already breaks down near the stagnation point on the back of the blade. According to figure 10 the interference fringes at a certain distance from the blade's surface disappear completely after a bend. That the area actually indicates the dead air behind a separation zone is plainly seen in figure 2J. The bend of the interference fringes at the back of the blade somewhat down- stream from the support wire breaks away from the surface of the blade . This indicates the surface of discontinuity which always emanates from an area of separation of flow. Such a surface of discontinuity dissolves in vortices. This is manifested in the washed out fringes as soon as the vortex frequency is sufficiently great. This is particularly plain from the separation at the face of the blade of figure 27. Larger vortices of correspondingly lower frequency are no longer 3een in the photograph j but they can be observed on the ground-glass plate by oscillation of the inter- ference fringes in the dead-air regions. The disappearance of the inter- ference line3 inside a narrow strip along the face of the blade in figure 10 is probably due to the deflected interference fringes being extremely thin at this point and not reproduced on the photograph. The Vigure3 10 to 30 indicate that the interference fringes in the boundary layer have a maximum deflection of eight interference fringe widths. Thus the maximum density difference in the boundary layer between ks: 3 blade wall and free flow is, by equation (12), Ap = 0.00997— ^r — • The ks s 2 m air density in the tests was about p = 0.12 —2- Since the change of m . . AT Ap state in the boundary layer can be regarded as isobaric, we get — = - -—, as it follows immediately from the equation of state for a small change of state. Therefore the maximum temperature difference inside the boundary layer is AT = 2^°. The displacement by eight widths, however, occurs only in very isolated cases. On the average it, and hence the temperature differences inside the boundary layer, are substantially less. 12 NACA TM No. 1209 dead- air zone at the "back of the "blade disappears again downstream. It is readily seen how, starting from the suspension wire, a new "boundary layer is formed in the accelerated flow. At the next greater flow angle, figure 11, however, the separation at the "back of the "blade haa already disappeared. But at the great angles of setting, figures Ik and 15, the flow "breaks down at the front side in the vicinity of the stagnation point. The flow adheres again farther downstream. In figure 13 a slight "breakaway is probably under way. The same phenomena recur for greater spacing ratios. At a/t = O.859, figures 16 to 21, a second region of separation is observed, at the "back of the "blade downstream from the end of the "blade passage. In figure 16 this second breakdown is not quite so conspicuous. A variation of the interference fringe right next to the "blade surface is noticed at the same place as in the subsequent photographs. Behind this region the flow remains separated. The result is an expansion of the wake "behind the "blade trailing edge. The exiting velocity direction, as shown "by the direction of the wake, is no longer exactly the same as the direction of the "blade trailing edge. The Hade grid is, at this spacing, no longer capable of completely deflecting the flow as far as the blade exit angle. The separation phenomena at a/t = 1.1^1, figures 22 to 27, are very pronounced. The wake behind the blades is very wide, as evidenced from the direction of the discontinuity surfaces arising from the regions of separation, and the direction of the outgoing flow is far from the blade exit angle. The blading is impractical for a turbine at this extremely large spacing ratio. According to the photographs the blading can be used up to a/t = O.859 and at flow angles ranging between 3^- and 62 . The slight separation on the back of the blpde at a/t = O.859 should not cause perceptible impair- ment of the efficiency, since the width ratio of the wake given by the rounding of the blade trailing edge to the width of the undisturbed air stream at the blade exit decreases with increasing spacing ratio. A sharp exit edge would reduce the width of the wake, so that the flow condi- tions would certainly be improved through the grid. Since there is a possibility that the separation phenomena at the blades are affected by the Reynolds number, the blading was investigated with a/t = O.859 at various flow velocities. Three photographs from this test series are represented in figures 28 to 30. It is seen that the separation is diminished at the higher speeds. If the Reynolds number is calculated using the axial component, 'w m , of the flow velocity and the blade chord t, its magnitude for figure 28 is 11800, for figure 29, 22300 and for figure 30, 35^00. The axial component was used for calcu- lating the Reynolds number because it can be precisely calculated on the actual turbine from the discharge volume and the flow cross sections. Gas or steam turbines operate with blades of from 1 to 3 centimeters chord and flow velocities in axial direction w m of from 100 to 300 meters per second NACA TM Wo- 1209 13 The gas - or vapor - temperatures range "between 1000 and 200 degrees and the pressures "between 10 and 1 atmosphere, except for maximum pressure turbines. This gives a Reynolds number ranging "between 10^ and 10°. In consequence the present test range lies approximately midway in the practical range. One condition for the applicability of the3e data is, of course, that the Mach number does not exceed the value 0.6 to 0.7, which usually is the case on gas turbines. VI. SUMMARY AND OUTLOOK To provide basic data for the design of turbine blading the flow through a blade grid of highly curved profiles was analyzed by the inter- ference method. The density of the air passing through the grid was determined from the records and the pressure distribution past the blades and the force produced by the pressures on the blade were obtained. Since the boundary layer of the flow at the blade surface is plainly visible on the interference records, any separation of flow is readily recognized. These separation phenomena were studied first. It is expedient to use comparatively large model blades in the tests. Limiting the Reynolds numbers to values usual for turbine operation, therefore, gives propor- tionally low airspeeds. One of the blade3 was slightly heated in order to make the boundary layer visible at these speeds also. The flow through a turbine blade grid with good (advantageous) blade profiles was examined at different spacing ratios and flow angles. The interference photographs of the flow through this grid as represented in figures 10 to 27 indicate that at the smallest flow angle (20°) the flow has already separated from the back surface of the blade just behind the inlet stagnation point. The flow separates likewise on the face of the blade at great flow angles (60 to 90 ). In both cases, however, the flow follows the blade closely again when the spacing ratio is small. At greater ratio (say from a/t = O.859 on), separation occurs again at the back of the blade near the trailing edge. Behind this region of sepa- ration, however, the flow no longer follows the blade surface. Thus the separation leads to expansion of the wake behind the blade and so certainly to a substantial impairment in efficiency. The described separation phenomena are affected by the Reynolds number, as was determined in a special test series (figs. 28 to 30). The tests are at present being extended to other blade forms, with the aim of developing suitable blade forms that ensure separation-free flow within a wide range of flow angles and spacing ratios. The tests are also to be extended to include blower blades. The flow losses can be determined by momentum measurements in the wakes behind the blades and the efficiency of the blading defined numerically. Since the temperature conditions in the boundary layer on the blade surface can be determined from the interference photographs, this method is particularly suitable Ik NACA TM No. 1209 for the study of the efficiency of the different methods for "blade cooling, The effect of the Mach number on the flow is to he investigated also. Translation "by J. Vanier, National Advisory Committee for Aeronautics. REFERENCES 1. Schardin, H. : Z. Instrumentenkde . Bd. 53, 1933, PP- 396-^03 and pp. 424- 43 6. 2. Zobel, Th. : Z. VDI, Bd. 8l, 1937, PP- 6l9"624. 3. Weinig, F.: Die Streaming um die Schaufeln von Turhomaschinen. (Leipzig), 1935, PP- 125-137- k. Ergehnisse der aerodynamischen Ver suchsanstalt , Gottingen, III. Lieferung, 1935, PP- 132-137- 5. Christian!, K. : Luf tfahrtforschung Bd. 2, 1928, pp. 91-100. 6. Hansen, G. : Z. techn. Physik. Bd. 12, 1931, pp. U36-HO. 7- Keller, C: Axialgehlase vom Standpunkt der Tragf liigeltheorie , Mitteilg. a.d. Inst, f . Aerodynamik. ETH (Zurich) 193^. 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