REPORT: 1368 SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIES COMPRESSOR BLADES AT LOW SPEEDS * By Jans C. Euupy, L. Joseeu Hunero, Joux R. Ban, and A, Riowanp Fours SUMMARY The performance of NACA 65-series compressor blade sections in cascade has been investigated syetematically in a low-speed eascade tunnel. Porous test-section vide walls and, Sor high-preseure-rise conditions, porous flecble end walls were ‘employed to establish conditions closely simulating two-dimen sional flow. Blade sections of design lift coeficients from 0 to 2.7 were tested over the usable angle-of-atiack range for tarious combinations of inletglow angle P of 80°, 45°, 60°, and 70°, and. solidity © of 0.50, 0.76, 1.00, 1.85, and 1.50. Design points were chosen on the basis of optimum high-speed operation. A suficient number of combinations were tested to permit interpolation and extrapolation of the data to all conditions within the usual range of application. The results of this investigation indicate a continuous varia tion of blade-section performance as the major cascade param- ters, lade eamber, inlet angle, and solidity were varied over the test range. Summary curces of the results have been prepared to enable compressor designers to select the proper blade camber and angle of attack when the compressor velocity diagram and desired solidity have been determined. At a few test conditions, an upper limit to the design lift co- effvient that could provide satisfactory performance was found. ‘These results provide information as to the mazimum value of the loading parameter, expressed a8 the product of solidity and section lift coeficient based on the vector mean velocity, that ean be used effectively in compressor design. Analysis of the trends indicated. that the common practice of employing « constant mazimum value of the loading parameter for all inlet angles and solidities fails to define the obserced performance of the compressor blades sludied in this investigation. Ar indes that the positive and negative limits of useful angle- of-attack range occurred when the section drag coeficient reached. twice the minimum ealue tons used to estimate the operating range of the compressor Blade sections studied. A broad oper- ating range for these sections was obserced, except for conditions of highest pressure rise across the cascade corresponding to high cambers at high inlet angles. These conditions are not typical of usual design practice and no dificulty should ordinarily be encountered in employing these Blade sections. In general, the obsereed performance of NACA 66-series compressor blades in cascade is considered to be very satisfactory. INTRODUCTION The design of an axial-flow compressor of high perform- ance involves three-dimensionel high-speed flow of com- ppressible viscous gases through successive rows of closely spaced blades. No adequate theoretical solution for this complete problem has yet appeared nor, from consideration of the complexity of the problem, does it seem likely that complete relationships will be established for some time. Various aspects of the problem have been treated theore cally, and the results of those studies aro quite usoful in design caleulations. All such studies, howover, havo been Dased on idealized flow, with effects of one or more such physical realities as compressibility, finite blado spacing, and viscosity neglected. Consideration of viscosity effects hhas been particularly difficult. Tt appears, thereforo, that in spito of advances in theoretical methods, theory ‘must ‘be supplemented by experimental data for some time (0 come. Some of the information required can be obtained only. by experiment in single-stage and multistage compressors. ‘Much of the information, however, can be obtained more easily by isolating tho effects of each parameter for detailed measurement. The effects of inlet angle, blede shape, angle of attack, and solidity on tho turing angle and drag produced can be studied by tests of compressor blades in ‘two-dimensional cescade tunnels. Cascade tests can pro- vide many basic data concorning the performance of com- pressors under widely varying conditions of operation with relative easo, rapidity, and low cost. A number of success- ful high-speed axiaLfiow compressors have been designed by using low-speed cascade data directly. A moro refined procedure, however, would uso cascade data, not as the final answer, but as a broad base from which to work out the threo-dimensfonal relations. ‘Date from a large number of two-dimensional cascade tests have been available in this end other. countries for *sopeades NAOA Techn Noe 216 dormarly NACA HO LALO) by T. Jeph Her, Tames 0, Beery, and Jobe en, 878 NAOA Tene Note 38 (one [AOR HAC Léa) by A, lad Fl, 0, nsm4 somo years. Although the cascade configurations used in these investigations were geometrically two dimensional, in no ease except that of the porous-wall cascade of reference 1 was the flow believed to be two dimensional. This situa- tion is ordinarily accepted on the grounds that the flow in the compressor is slso subject to three-dimensional end cffects. ‘That similar end conditions would exist in. ste- tionary cascades and rotating blade rows seems unlikely. As discussed in reference 1, there is evidence, however, thet the flow through typical axial compressor blades is more nearly like that in aerodynamically two-dimensional cas- cades than like that in cascades which are only geometri- cally two dimensional. Excellent correlation between porous-wall cascade and rotor-blade pressure-distribution ‘and turning-angle values is shown for the design conditions ‘of the compressor investigation reported in reference 2. ‘Pho proper basic approach to the compressor design prob- lem, therefore, would seem to be to separate the two- dimensional effects from tho threo-dimonsional. This ap- proach should also aid in the evaluation of the separete affects of tip clearance and secondary fiow in axial com- pressors. Therefore, systematic series of low-speed cascade tests of the NACA 65-series compressor blade sections were made by means of the porous-wall technique to insure two dimensionality of the flow. ‘These results and an analysis of the data aro presented, as well as sum- marized data in the form of carpet plots for use in specific design problems. syMBoLs a mean-line loading designation ‘ blade height or span, feet © Dlade chord, feet on seotion drag coefficient o resultant-force coefficient ¢ section lift coefficient fue camber, expressed as design lift coefficient of isolated airfoil - ew section normel-fores coefficient, evar section normal-force coefficient. obtained by cal- culation of momentum and pressure changes across blade row exe section normal-forco coefficient obtained « by integration of blade-surface pressure distribution & wake momentum difference coefficient F force on blades, pounds Fy foreo on blades due to momentum change through blade row, pounds F, oreo on blades due to pressure change through Dlade row, pounds Fz foro on blades due to momentum difference in ‘wake, pounds ® tangential spacing between blades, feet Ud ratio of section lift to section drag Mo Mach number P total pressure, pounds per square foot P statie pressure, pounds per square foot ‘Ap _ static pressure rise across cascade, pounds per ‘square foot ‘REPORT 1368—NATIONAL ADVISORY COMAUTTER FOR AERONAUTICS dynamic presure, pounds per square foote pla, nondimensional static-prescuresiso peramter Reynolds number based on blade chord pressure coefficient (Fe) rotor-blade rotational speed, fest per second flow velocity, feet per second flow velocity relative to blades, feot per second chordwise distance from blade leading edge, feot sealed value substituted for 6, 0+50(cr.—0.4) perpendicular distance from blade chord line, fect angle between flow direction and blade chord, degrees tangle between flow direction and a perpendicular to the cascade axis, degroes flow turing angle, degrees (8,~8.) angle between resultant-foree direction and tan- gential direction, degrees mass density of flow, slugs per cubic foot solidity, e/g subscripts: ‘component in axial direction design value, excopt in drag coefficient local referred to veotor-mean velocity, Wn flow outside wake component in tangential direction upstream of blade row downstream of blade row APPARATUS, TEST PROGRAM, AND PROCEDURE DESCRIPTION OF TEST EQUIPMENT ‘The test facility used in this investigation was the Langley S-inch low-speed, porous-wall cascade tunnel described in reference 1 and shown in figures 1 and 2. During the course of this program some further improvements wore required to establish proper testing conditions at higher pressure rise conditions. In particular, the boundary-layer buildup behind the slot on the convex flexible end wall with high pressure rise cascades was sufficient to cause separation and destroy simulation of the infinite cascade even though the blade flow was not separated. ‘This condition was corrected by replacing the end wall with a porous flexible wall and suction chamber. In addition tho large difference in flow Jongth from the entrance cone to the side-wall slots between the tunnel ends at the higher inlet air angles gavo quite different boundary-layer thickness along the length of tho side-wall slots and made uniform flow entering tho test section difficult to obtain. This condition was improved by making the changeable side plates porous and drawing small emount of air through them. ‘The concave flexible, ‘end wall was made porous to provide a further control of flow conditions through the test section. ‘The porous material found to be most satisfactory is com~ mercial woven monel filter cloth. ‘This cloth is available in various meshes in widths up to 36 inches and can bo calendered at the factory to reduce porosity and improve surface smoothness. ‘The combination found most suitable for the present work wes e Dutch twill double weave of 30 Gee 2s w REE GG & HBO vueegeaeSYSTEMATIC TWO-DIMENSIONAL CASCADE THSTS OF NACA 65-SERIES COMPRESSOR BLADES AT LOW SPEEDS by 250 mesh with warp wire diameter of 0.010 inch and fill wiro diameter of 0.008 inch. ‘The original thickness of about 0.026 inch was reduced to 0.018 inch by calendering. ‘Tho resulting material has the porosity characteristics shown in figure 3. ‘The primary advantages of this material over others tried previously are its uniformity, flexibility, strength, surface smoothness, and relatively low cost. ‘Tho blade family used in this investigation is formed by combining basic thickness form with cambered mean lines. ‘The basic thickness form used is the NACA 65(216)~ 010 thickness form with the ordinates increased by 0.0015, times tho chordwise stations to provide slightly increased thickness toward the trailing edge. This thickness form hhas been designated the 65-010 blower blade section in references 3 and 4. It wos not derived for 10-percent thickness but was sealed down from tho NACA 65,2-016 airfoil given in reference 8. As discussed in reference 5, the scaling procedure gives the best results when it is restricted to maximum thickness changes of afew percent. Subsequent to the enscade studies of reference 3, in which the NACA 65-010 blower blade was first introduced, the basic thickness form for the NACA 65-010 airfoil section was derived and ineluded in reference 5. ‘The ordinates for the basic thick- ness form of tho blower blade in reference 3 differ slightly from the airfoil ordinates in reference & but are considered interchangeable within the accuracy of the results reported herein, Ordinates for both the sealed thickness form (ref. 3) and derived thickness form (ref. 5) are given in table I. TABLET ORDINATES FOR NACA 65-010 BASIC THICKNESS FORMS [Stations and ondinates in pereent chord) Ondinates, y Station, | 65@216)-010 | Derived fitfoil ome | 05-010 bined wath | ‘sito peoooise 0 ° 152 me 300 a3 rh 1189 ait iar 2m au 2709 pai ain 2040 ae 3.068 221s 4143 i500 50a Sau 4780 4 ose Som ‘Base £000 3020 oa 4570 aan S50 £530 cia rar Sear ‘082 ons ais Rast 2 bea Lar Los 136 1385 49 B10 Tat 308 150 60 687 715 ‘The basic meen line used is the a=1.0 mean line given on page 97 of reference 5. ‘The amount of camber is expressed in reference 5 as design lift edeffcient cx» for the isolated air- foil, and that system has been retained. Ordinates and slopes, for the a=1.0 mean line aro given in table II for c,.—1. Both ordinates and slopes are sealed directly to obtain other cambers. Cambered blade sections aro obtained by applying the thickness perpendicular to the mean line at stations laid out along the chord line. The blade sections tested are shown in figure 4. In the designation tho camber x is given by the first number after the dash in tenths, For example, the NACA 65-810 and NACA 65-(12)10 Blade sections are cambered for c=0.8 and 2, respectively. [TRET PROGRAM AND PROCEDURE ‘Test program—The test program was planned to provide sofficient information to satisfy any conventional vector diagram of the type shown in figure 5. ‘Tests of seven- blade cascades were run with various combinations of inlet air angle 6; of 30°, 45°, 60°, and 70°, solidity « of 0.50, 0.75, 1.00, 1.25, and 1.50, and eumbers from cx, of 0 to 2.7 over the useful anglo-of-attack range. ‘The most complete test sories wer run at solidities of 1.00 and 1.50; suflicient tests, were made at the other solidities to guide interpolation and extrapolation, ‘Tho combinations of 6, 0, and blade sec- tion which were tested are tebulated in table TIT. The camber range covered at solidties of 1.00 and 1.50 was determined by one of two limitations. At the higher inlet angles. progressively higher cambers were used until tho limit loading had been reached, that is, until the design condition coincided with stall; at lower inlet angles, how- ever, design turning angle exceeded inlet anglo before the limit loading had beon reached and the tests were terminated ‘TABLE TL ORDINATSS FOR THE NACA a=1.0 MEAN LINE, ey=10 [Stations and ordinates in percent chord] ni ome 280 | 593150,‘TABLE II CASCADE COMBINATIONS TESTED ow Hilde] Bib eign] Sins | ee Eine | SR ae [oe apse eae | ese | ae | Bb io| Hi ng| Hite | Wile | ER) Ae RR | CURE at at em | Se. Se vs Ee| Ei | ae #8) HB8| Bb eg ae at as i ie | Hae | eae ve | Bib) ie] Ete) Bib eine] eepe) tine) BiB SSE] RH Su ee) ae SHER | SEAN “No design point was obtained for this combination, there, Limit conditions were attained at 670°, «1.00, 41.25, and 1.50, and at ,=60°, o=1.00 and 1.50. ‘Test procedure.—It was shown in reference 1 that two- dimensional flow can be achieved by controlling the removal of boundary-layer air through porous test-section side walls so that the downstream static pressure equals the ideal value, corresponding to the turning angle, corrected for the Dlocking effect of the wake. This criterion wag accordingly used in these tests. Tn addition, the flexiblo end-wvall shapes, and suetion-flow quantitios were adjusted to obtain uniform ‘upstream flow direction end wall static pressures, criteria, of tivo-dimensional flow simulating an infinite cascade. ‘This proceduro necessitated an approximate meesurement, of turning angle and wake size and an estimate of the correct iso before the final settings could bo made. Initially, this system requived some cut-and-try procedure but after the initial tests at each combination of f and «a chart similar to figure 5 of reference 1 could be drawn to assist in estimating the pressure rise. An experienced operator could make the required estimates and settings very quickly by this procedure. Spot calculations of the correct pressure rises were made after completion of tests to check the accuracy of the values used. ‘Tests wero made at each cascade combination shown in tablo IIT over a range of angles of attack at intervals of 2° to 3°, In general, the tests covered the interval from nega- tive to positive stall, where stall was determined by a large {ncreaso in wake size. ‘The principal exeoptions occurred for Tow cambered blades where negative stall would have REPORT 1368—NATIONAL ADVISORY COMMITTEE POR ABRONAUTICS occurred at negative turning. It was found that the small wall boundary-layer buildup for negative turning angles and hence negative pressure rises would heve required a less porous material than that normally used, to avoid excess air removal while maintaining sufficient suction pressure diff ential to avoid local reverse flow through the porous material, Te was not deemed worthwhile to change the porous material to obtein data in this relatively uninteresting range. For the NACA 65-010 section at i=80°, however, the diffieulty persisted well above 0° turning, and this combination was tested with both porous and solid walls. Tho tests wore entirely within a speed rango considered incompressible. ‘Tho Bulk of the tests at solidities of 1.00 and 1,50 wore run at an entering velocity of 95 feot per second. For tho usual 5-inch blade chord, the Reynolds number was 245,000. Some information near the design point was obtained at higher effective Reynolds number for ‘most cascade combinations by adding roughness to the blade Teading edges in the form of ¥einch-wide strips of masking tape draped around the leading edges from wall to wall. In addition, some tests near design conditions were run at speed of 135 feet per second and a Reynolds number of 346,000 with and without roughness. ‘Two cascade com- binations were tested at design conditions over a rango of Reynolds numbers from 160,000 to 470,000 to assist in esti- mating performance at Reynolds numbers other than the usual test value. In order to provide further information on sealo effects, two cascade combinations wero tested through the range with leading-edge roughness at the standard Reynolds number and in the smooth conditions at a Reynolds number of 445,000. For solidities of 0.50 and 0.76 the tunnel could not accommodate seven blades of 5-inch chord; the blade chord was reduced to 2.5 inches and the Reynolds num- ber to approximately 200,000 for those tests. Tests with roughness were made near the design point for solidities below 1.00, but Reynolds numbers higher than 200,000 could not be obtained with the existing equipment Test measurements.—The blade pressure distribution was measured at the midspan position of the central airfoil at cach angle of attack. In addition, surveys of wako total- pressure loss and turning angle were made downstream of the cascade. ‘The total-pressure surveys were made with a non- integrating multitube rake approximately 1 chord down- stream of the blade trailing edges. ‘Turning anglo surveys ‘wore made by the “null method” with a claw-type yaw hond; since the yaw dovieo was mounted on track at the rear of the tunnel the distance from tho blades varied from about 1 to 8 chords in tho flow direetion depending upon the inlet~ and turning-angle combination. Flow discharge angle readings were taken at a number of points downstream of several blade passages along the tunnel center line. These readings were averaged to obtain the final value, Sinco the angle readings in the wake deviated several degrees from the average reading, and the direction of tho deviation varied consistently with the direction of the total pressure gradient, tho accuracy of readings in the wake was questioned. ‘There- fore, the values obtained when the wake readings wore included and excluded in the averaging process wore com- pared for a number of tests. ‘The resulting turning-angleSYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SHRIRS COMPRESSOR BLADES AT LOW SPREDS curves compared very wel, but scatter was considerably less ‘when only the readings outside the wake were used to obtain the turning-angle value. This latter procedure has been ‘adopted as the standard method of measuring the flow dis- chargo direction. Static pressure and upstream flow angles wre measured approximately 1 chord upstream of the blade row. ‘Total pressure was measured in the settling chamber. Pressures were obtained by orifices with pressure leads to manometer boards, Angles of flow were again obtained by use of a claw-type yaw head by the null method. Csloulations —Pressure distribution and wake-survey data were recorded and force values calculated nondimensionally 3 coefficients based numerically on the upstream dynemic pressure q1. This choice was dictated partly for convenience in reducing tho data (standardization of g, permits use of manometer scales which give nondimensional values directly) ‘and partly because information based on entering flow was considered most convenient for design use, particularly when critical speed is important. All forces due to pressure and momentum changes across the blade row were summed to obtain the resultant blade {oree coefficient ér.. In this process the wake total-pressure deficit was converted to an integrated momentum difference by the method given for the drag caleulation in the appendix of reference 5. ‘This wake momentum difference, expressed nondimensionally, is designated tho wake coellicient cus; it represents the momentum difference between the wake and the stream outside the wake, and is based on q, numerically. ‘The wake coeficient is not considered to be a true drag cooficient, but is used merely for convenience jn assessing tho contribution of the wake in the summation of forces. ‘The resultant-force coefficient: was resolved into compo- nents perpendicular and parallel to the vector mean velocity Wr (Gee fig. 6) to obtain the lift coefficient cx, and the drag coefficient a1, respectively. ‘The mean velocity was cal- culated as the vector average of tho velocities far upstream and far downsiream. The velocity far downstream was obtained from measurements just behind the blades by using a momentum-weighted average of the velocity just behind the blades, ‘This rather detailed mothod was found neces- sary to give consistent drag values, Since the resultant force is very nearly perpendicular to the mean velocity, the value of the component parallel to the mean velocity is quite sensitive to mean-velocity direction. Attempts at using the downstream velocity outside tho wake for averaging rather than the momentum-average velocity gave very erratic drag results and indicated that mean velocity directions obtained in that manner were not reliable, Tn addition to the lift and dvag, the section normal-force coefficient ey xr. Was obtained by computing the component of the resultant-force coefficient perpendicular to the blade chord line. ‘This normal-forco coefficient was compared with the normal-force coefficient év.ra obtained by integration of the blade surface pressure distribution os a check on the accuracy of individual tests. A detailed derivation of the method of calculating the foree coefficients is given in appendix A. Accuracy of results.—In general tho tuming-angle values measured are believed to be accurate within +£0.5° near the design values. ‘The correlation procedure used is believed nT to have improved further the accuraey of the design values in the final results. For tests far from design, that is, near positive or negative stall, the accuracy was reduced some- what. Tn addition, at an inlet angle of 70° with sections of zero camber, satisfactory measurements were very difficult to obtain and the accuracy was reduced. ‘As noted in the section describing calculation methods, the Dlade normal-force coefficient cy a, calculated from pressure- rise and momentum considerations was compared with the normal-force coefficient cy.p. obtained from the pressure distribution as a check on the overall accuracy of individual tests. Since these values would be affected by error in tuming-angle, surface pressure or wake-survey readings, or by failure to achieve two dimensionality of the flow, this ‘comparison is a check on the overall acceptability of the results, A difference of 5 percent between the two normal- force couffcients was set as the outside limit for acceptance of individual tests for lift coefficients above 0.2; below lift cooflicients of 0.2 a direct numerical comparison was made using a limit of +£0.01. Tho agreement was well within ‘the 5-pereent limit for most of the tests as originally run, and only a few conditions had to be repeated. ‘The accuracy of the lift coeficients is directly comparable to that of the normal-force coefficients. Tho accuracy of wake-cocfficient and dreg-coeflicient values will be discussed Inter under ‘Reynolds number effects PRESENTATION OF RESULTS Detailed blade-performance data for all cascade combina- tions tested are presented in the form of surface-pressure distributions and blede-section-characteristic plots in figures, 6 to 84. The represuntative pressure distributions presented havo been selected to illustrate the variation through the angle-of-attack range for each combination, The section characteristics presented through the angle-of-attack range are turning angle 6, lift coellicient cis, wake coefficient ew.1, drag coefficient cg,, and liftdrag ratio Ud. ‘Tho effects of ‘changes in Reynolds number end blede-surface condition on section characteristics are given in figure 85. ‘Trends of section operating range, in terms of angle-of- attack range, with eamber for the four inlet angles of the tests are presented in figure 86. Variation of ideal and actual dynamic-pressure ratio across tho cascade with turning anglo and inlet angle is presented in figure 87. Figure 88 gives the relation between inlet dynamic pressure and mean dynamic pressure for convenience in converting coefficients from one reference velocity to the other. Limi loading information is summarized in figure 89. Comparison ‘of the present porous-wall-cascade tuming-angle data with those of the solid-wall cascade of reference 3 is made in figure 90 for a typical inlet angle and solidity combination ‘The information which is most useful for choosing the blade sections to fulfill compressor design vector diagrams is summarized in figures 91 to 111. The variation of turning angle with angle of attack for the blade sections tested is, presented for one combination of inlet angle and solidity in each of the figures 91 to 106. ‘Trends of the slopes of the curves of turning angle against angle of attack near design are given in figure 107. Figures 108 to 111 are design and correlation charts; the variation of design turning angle and718 design anglo of attack with the parameters, eamber, inlet angle, and solidity, is presented for several combinations of the parameters so that interpolation to the conditions of a design velocity dingram is relatively easy. ‘The selection of the blade section and the blade setting required to fulfil a given velocity diagram at design condi- tions requires several interpolations and cross plots of the date presented in figures 6 to 84, In order to facilitate these interpolations, a carpet plotting technique described in reference 6 has been used to correlate the date, This technique, for whieh details are explained in appendix B, permits a function of several variables to be presented on @ single graph and such a graph lends itself readily to interpola- tions for intermediate values of the variables plotted. Carpet plots summarizing the present date are presented in figures 112 to 116, and an example illustrating their use is included in appendix B. DISCUSSION DESIGN CONDITIONS ‘The values and shape of the blade-surface pressure distribu- tion aro important eriteria for predicting the conditions of best operation at high Mach numbers. Velocity peaks oc- curring on either surface in low-speed tests would be ac- centuated at high speeds, and supersonic velocities with at- tendant shock losses would occur at relatively low entering ‘Mech numbers. ‘The selection of the angle of attack desig- nated as “design” for each combination of inlet angle, solidity, and camber is based on the premise that the blade section Will be required to operate at Mach numbers near tho critical value. ‘The trend of pressure-distribution shape over the anglo-of-attack range was examined for each cascade combination, and the angle for which no velocity peaks occurred on either surfece was selected as being optimum from the standpoint of high-speed usage. In general, the design angle so selected is near the middlo of tho low-drag. range thus indicating efficient section operation for angles, a fow degrees higher or lower than the design condition. ‘The choices of design angle of attack are indicated by an arrow on the bladesection-characteristic plots of figures 6 to 84. ‘The design angles are also shown by cross bars on the turning- angle summary curves in figures 91 to 108. Correlation of the design angles of attack and design turn- ing angles over the rango of camber, solidity, and inlet angle is given by figures 108 to 111 in 2 manner convenient for design use. ‘The correlation is excellent; smooth curves re- sult when any two of the three parameters aro used as inde- pendent variables. ‘The section-characteristic curves of figures 6 to 84 indicate that, in general, the design points chosen do not give maxi- ‘mum lift-drag ratios for low- and medium-speed operation. For designs which will not operate near eritieal speed, there- fore, higher efficiency could be obtained by using angles of attack several degrees higher than the design points pre- sented. ‘This procedure must be used with caution, however, at the higher camber and inlet-angle combinations since the section operating range becomes quite narrow for combina- tions of highest camber and inlet angle corresponding to REPORT 1308—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS tho highest values of Apia. It is recommended tht the individual pressuro distributions and soction-characteristio, ccurves bo examined before departure from the specified design points is made. Pressure distribution and boundary leyers.—Vor many of the tests at angles of attack near and below design, there is evidence that a region of laminar separation of the boundar layer flow occurred on the convex blade surface; this sopa- rated boundary layer then became turbulent and reattached to the blade surface as a relatively thick turbulent boundary layer. The mechanism of such a flow sequence is deseribed for the isolated airfoil in reference 7. ‘Tho laminar separation is indicated by a relatively flat region in tho pressure distri- bution on the convex surface and tho turbulent reattachment is characterized by erapid pressure recovery just downstream of the separated region. ‘This low pattorn can be seen clearly ‘in many of tho figures but is particularly evident in figures 42 (a), 42 (0), 65 (b) to 56 (d), and 60 (a) to 66 (@). For some tests, Inminar separation appeared to occur on the concavesurfaceaswell. This is noticeable in figures 42 (b), 42 (@), 42 (¢), 66 (6), and 66 (€). ‘The extent of laminar boundary-layer flow which occurs on ‘an airfoil surface is affected by Roynolds number, stream turbulence level, airfoil surface condition, and surface pros- sure gradient. Increases in Reynolds number, stream turbu- once, and surface roughness would promote earlier transitio Qualitatively a gradient of decreasing surfaco pressure would be required to maintain laminar flow if the Reynolds number, stream turbulence, surface roughness, or the combination of these, which might be referred to as “effective Reynolds number,” were high enough to favor transition. At the turbulence level of the S-inch cascade tunnel, however, aminar flow and laminer seperation on the convex surface persisted to Reynolds numbers up to 248,000 even when the surface pressure gradient was slightly unfavorable. ‘Tho addition of leading-edge roughness, as described in the section “Test Procedure” reduced the extent of the laminar separa~ tion region, but did not eliminate it in somo eases. In view of the thick boundary liyer which results from laminar separation and reattachment, it appears thet the minimum final boundary-layer thickness and section drag coefficient ‘would result if the Reynolds number and turbuleneo values were such as to cause transition before laminar separation occurred. Use of leading-edge roughness to reduce an ex- tended laminar separation region would probably result in a thinner final boundary layer than that for the smooth blade at tho same Reynolds number but would probably result in a thicker boundary layer than that for the smooth blade at hhigh Reynolds number. A thick turbulent boundary layer would be expected to promote turbulent separation near the trailing edge of compressor blades which produce a signifieant ‘pressure rise. Wake coefficient and drag coefiicient—As noted pro- viously, the wake coefficient o., expresses the momentum difference between the wake flow and the downstream flow outside the wake in manner convenient for use in summing blade forces. ‘The wake cosffcient is, of course, directlySYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-ERIES COMPRESSOR BLADES AT LOW SPEEDS dependent upon boundary-layer thickness and shape, and changes in the boundary layor with changes in effective Reynolds number are refleoted in the wake-coefficient values. Furthermore, if the effective Reynolds number is neer the condition where laminar separation may or may not occur, the change in surface pressure gradient with change in angle of attack would control the presence and extent of Jeminar separation on either blade surface. Obviously erratic varia tions in the value of the wake coefficients would result under those circumstences. ‘The blade-section-characteristic curves of figures 6 to 84 show that in most cases the wake-coeflicient values woro irregular as the angle of attack was varied in the region near design at the usual test Reynolds number of 246,000. With either higher Reynolds number or leeding- edge roughness, or both, the rapid local pressure recovery associnted with boundary-layer reattachment was less evi- dent, in tho surface pressure distributions and the wake cocflicient usually was reduced. For a few cases, notably those of figures 34 (), 35 (g), 68 (g), and 8¢ (¢), leading-edgo roughness increased the wake coefficient, however; in those cases the roughness apparently produced a more severe turbulent boundary leyer than laminar separation and re- attachment did. ‘The trend of deag coefficient c41, defined as the component of resultant force parallel to the mean velocity, was similar to that of wake coefficient. ‘The drag curves were quite irregular near design angle of attack and the values measured varied as much as 30 pereent with Reynolds number and roughness. Obviously the values of both drag coefficient and lift-drag ratio near design are not sufficiently reliable to use direetly in design analysis. ‘These values should be of some use for comparison purposes, however. ‘The large drag rise associated with positive and negative stall should be relatively insensitive to Reynolds number effects, because tho pressure gradients on the critical surfaces are then un- favorable to laminar flow and therefore should be useful for determining offective operating range. ‘The trend of drag coefficient with Reynolds number nest the design condition for the NACA 65-(12)10 blade section at By of 60°, ¢ of 1.00, and A, of 45°, « of 1.50 shown in figure 85 (a) sorves to indiente the magnitude of the Reynolds number effect. Increasing the stream turbulence by the use of a }inch-mesh sereen upstream of the test section lowered the drag coefficients at low Reynolds number, and reduced the Reynolds number at which the drag coefficients become essentially constant with Reynolds number. The compari son of eg, values through the angle-of-attack range for the samo easeade combinations at two Reynolds numbers in figures 85 (b) and 85 (0) gives some further indicetion of Reynolds number effect. For R of 445,000, tho drag cooffi- cients are lower and the curves are smoother than for 7 of, 245,000. ‘The addition of leading-edge roughness in figure 86 (b) smoothed the drag curve but did not give the same deorense in drag that the high R did. ‘There appears to be some effect on the angle of attack at which the drag rises rapidly in figure 85 (c) but since the effect: was not the same in figure 85 (b) no conclusions can be drawn. ‘Turning angle and lift—Figure 85 (x) shows that the effect of Reynolds number on turning angle near design ay is 20507-0048 79 almost insignificant for R between 220,000 and 470,000. ‘This is borne out by the fect that throughout figures 6 to 84 changes in @ with Reynolds number and roughness were, in general, within the limits of measuring accuracy. Below R of 220,000 a decrease of design turning angle ean be expected. Reynolds number appears to have some effect on tuming angle near stall in figure 85 (c), but again the effect has not been definitely established. Tt can bo concluded that, the design turning angles presented are correct for R above 220,000, but that the effect of near stall is unknown. ‘Laminer separation hed no appreciable effect on the measured lift. ‘The lift-coofficient values for a given test, agreed well at low and high Reynolds numbers and with and without roughness. ‘The normal-force coefficients ob- tained by integration of the pressure distributions also changed very little with changes in Reynolds number and roughness, In order to estimate the useful operating angle-of-attack range of the various sections at the several solidity and inlet angle conditions tested, Howell's index of twice tho mini- ‘mum drag (ref. 8) was used to select the upper and lower limits of angle of attack. As discussed previously in the section concerning Reynolds number effects, the accuracy of the measured values of drag coefficient near design angle of attack suffered due to laminar-flow separation. The minimum value of drag coefficient could not be determined exactly end an approximate value was used to determine the operating range. For most of the test configurations, the drag coefficient changed rapidly with angle of attack near the ends of the useful range, so an error in the value of mini- ‘mum drag used would have only small effect on tho operat ing range value. Some scatter in the results was evident, however. No significant effect of solidity was observed. Most values at constant camber and inlet angle fell within the scatter of the points. A tendency for the range to increase slightly as tho solidity was increased seas detectable at p,=45°, but this was not evident for other inlet angles. ‘The results plotted in figure 86 indicate that the major determinant of the oper- ating range is inlet angle. As the inlet angle is increased, the usable range of angle of attack is decreased, with greater changes indicated for angles above design than for angles below design. ‘The camber of the section affects the oper- ‘ating range in the following manner for angles of attack above design: at an inlet angle of 30°, tho range increased with increasing camber; at inlet angles of 45°, 60°, and 70°, the opposite trend occured. For values of a less than design, little change in range with camber was indicated for :=30°; at higher inlet angles, the range decreased as the section camber incressed. With high entering velocities, the section operating range would be reduced because of a more rapid increase of drag ‘at angles of attack well above or below design. Further, the comparison between sections of different camber, at constant inlet angle and solidity, would be altered es the flow velocities relative to the blado surfaces exceed the local velocity of sound.720 ‘Tho ideal, nondimensional pressure rise Ap/gy across a. two- dimensional cascade is specified when the inlet angle and turning angle aro known, since the ratio of the flow areas determines the pressure rise. Since the mass flow is con- stant, the actual pressure riso is less than the ideal because of the “blocking” effect of the wake on the downstream flow area. For given inlet and turning angles, the blocking effect ‘would be more severe for higher solidity, tince the unaffected flow area is reduced. For incompressible flow the nondi- ‘mensional pressure rise is equal to one minus the dynamic- pressure ratio, thet ia, S21, ‘Tho actual dynamic. a 4% pressure ratio becomes higher than the ideal because of the wake blocking effect. ‘The ideal dynamic-pressure ratios, and the actual ratios at design turning angles for two solid~ ities, aro summarized in figure 87 for the range of inlet and turning angles of the tests. ‘The dynamie-pressure ratios for individual tests aro given by the short bars at the 100- percent points of the pressure-distribution plots in figures 6 to 84, Wake blocking effeots would be changed by the same Reynolds number and roughness factors which change the wake coefficient; however, the percentage change in dynamic- pressure ratio would be small. Tnformation on the maximum Ioading which can be achieved in a compressor blade row is important in tho design of high performance axial-low compressors. As noted viously, the high pressure riso associated with large turn- ing at high inlet angles promotes turbulent separation so that at inlet angles of 60° and 70° the stall angle of attack moved progressively closer to tho design value with inereas- ing section camber. ‘The limit turning is reached when the ‘maximum turning angle is no greater than design turning angle. The practical limit would be somewhat lower to sive a reasonable operating range. Approximate limit turning was reached at 6, of 60°, ¢ of 1,00 and 1.50, and at 6, of 70°, ¢ of 1.00, 1.25, and 1. Information from those tests is given in terms of s commonly ‘used loading parameter, o¢im) in figure 89. Both the actual test values of the parameter, and the ideal values calculated using the test inlet and turning angles are presented. Nota that the lift coefficient is here based, numerically, on the mean velocity, to conform to the usual form of the parameter. Arbitrarily chosen constant values of cin have often been used os maximum allowable values in design analyses. ‘Tho fallacy of using any constant value as a limit is clesely shown in figure 89; the true limiting value increases with increasing solidity and decreases with increasing inlet angle. “Since no limits were reached for inlet angles of 45° and 30°, itis clear that the limitation bas very little significance there except, pethaps, at very low solidities. ‘The phenomenon is not yet ‘well enough understood to permit the choice of parameter which could define the overall limitation as e single value. REPORT 1363—NATIONAL ADVISORY COMMUTER FOR AERONAUTICS COMPARISON WITH SOLID-WALL CASCADE DATA ‘The comparison between pressure-distribution and turn- ing-angle data for a solid-wall cascade tunnel and for the present porous-wall cascade tunnel is given in vefereneo 1 for the NACA 65-(12)10 blade section at 6; of 60° and o of 1.00. ‘The comparison has been extended in figure 90 to include turning-angle date for all the cambers reported for 6; of 60° and ¢ of 1.00 in reference 3. ‘The turaing-angle curves com- pare fairly well for cambers up to ex,» of 0.8, but for the airfoils of higher camber the deta of reference 3 doviate significantly from the present results. Comparisons at other conditions would show similar trends. Summaries of the relationships between turning angle and angle of attack through the camber range are given for each inlet angle and solidity in figures 91 to 106. ‘The variations are quite consistent for most of the range. Some ineonsist- ‘ney in the shapo of the curves at stall isa result of reduced. accuracy of measurement there. For combinations giving moderate pressure rises straight-line relationships aro indi- catod for considerable portions of tho curves. For the high- est pressure rises, however, no definite straight-line relation- ships exist. ‘The variation of the slopes near design is given in figure 107 to assist in estimating relationships at conditions other than those tested. ‘These slopes aro average slopes for tho camber range, and do not apply for the highest eambers. ‘They must be used with particular caution for inlet angles near 70°, since very nazrow straight-line regions axe prevalent there. ‘The usuel procedure in blade-section selection is to deter mine the camber ei, which is required for a given design velocity diagram at a selected solidity. Figure 112 gives carpet plots of ‘the data at five solidity conditions. ‘The carpet plots indicate the variation of camber ¢,», at design angle of attack «,, with required values of inlet-air angle 8, and design turning angle 0. Ench carpet plot is spaced from the next by a number of grid units proportional to the differ- ence in solidities. Since design angle of attack is independent of inlet-air angle, it is possible to present a carpet plot (Gg. 118) showing design angle of attack ay a3 a funtion of camber ¢,, and solidity «. ‘The tesis were mado at fixed inlot-air angles with the angl of-attack variation produced by changing the blade setting. Although date of this type facilitated the determination of design conditions for the various combinations of inlet-nir angle, solidity, and camber, it does not lend itself casily, as presented, to obtaining off-design performance of a blade section as regards operation of this section in a compressor in which blade setting is fixed and the angle of attack varied by changing the inlet-air angle. However, if.tho date. are plotted as an off-design carpet, itis a simple matter to draw in curves of constant blade setting and thus to predict theSYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA O5-SERIES COMPRESSOR BLADES AT LOW SPEEDS variation in turning angle with angle of attack for all inter~ mediate conditions of solidity, eamber, and blade setting. Such off-design carpet plots showing turning angle as a fune- mn of solidity, inlet-air angle, angle of attack, and camber aro presented in figure 114. Off-design data are presonted for the following sections: NACA 65-(4)10 NACA 65-(8)10 NACA 65-(12)10 : NACA 65-(15)10 NACA 65-(18)10 Pressure rise as a percentage of inlet dynamic pressure play, called pressure-rise coefficient, has been used a3 2 cascade londing-limit parameter. It is known thet cascade losses increnso rapidly above certain limiting values of Aplq. However, in view of the physical meaning of the pressure-tise coefficient, particularly in reference to the inner nd outer casings, it is considered to be a useful parameter. If the inlet-air angle, the turning angle, and the entering Mach number aro known, calculations of an isentropic pressure-tise coefficient. is possible, provided some relation- ship is assumed between the entering and leaving stream- tube areas. Therefore, tivo carpet plots were made in order to show the variation of the pressure-rise coefficient. with inlet-nir angle 6, turning angle @, and entering Mach number My. The first of these plots (fg. 116 (a)) was calculated by ‘assuming constant stream-tube area; the second (fig. 116 (b)) ‘was calculated by assuming that the stream-tube area varied so that constant axial velocity would be produced across the blade passage. Pressure-rise coefficients obtained from 721 these two plots very often bracket the value associated with the actual three-dimensional condition being examined. SUMMARY OF RESULTS ‘Tho systematic investigation of NACA 65-series compressor blade sections in a low-speed cascade tunnel has provided design data for all conditions within the usual range of application. ‘The results of this investigation indicate a continuous variation of blade-section performance as the important, cascade parameters blade camber, inlet angle, ‘and solidity are varied over the useful range. Summary curves have been prepared to facilitate selection of blade sections and sottings for compressor-design velocity diagrams for optimum high-speed operation. ‘Upper limits for the loading parameter aja have been established for some conditions, and the invalidity of using ‘constant value of the parameter has been shown. ‘The variation of the useful section operating range with camber, inlet angle, end solidity has been shown. ‘The ‘operating range was found to be broad except for the highest pressure-rise conditions. Compressor-blade cascade data have been presented in the form of design carpet plots, which greatly facilitate the selection of compressor blade sections required to fulfill ‘velocity diagrams. Plots of this type also are shown to increase greatly the usefulness of available cascade data by iding a simple method of obtaining the off-design varia in turning angle with angle of attack. Lanouey Awnonavricat Lanonavony, ‘Namtowat, Apvisony Cosnurrse Fon AERONAUTICS, Laxauer Frau, Va., January 31, 1958.APPENDIX A CALCULATION OF BLADE FORCE COEFFICIENTS, ‘The two-dimensional resultant force on a blade in cascade i the veetor sum of all tho pressure and momentum forces exerted by the fluid. At any appreciable distance behind the blade row the static pressure is constant along a line parallel to the blade row, since any prior pressure gradients, ‘would have been converted to momentum changes. On tho assumption that pressure foree acting inthe upstream ditec- tion is positive, P= Odio @ sum the momentum forees in the exial end tangential direc- tions. Assume that the axial momentum forces are positive if the force on the blade is in the upstream direction and that the tangential momentum forces are positive if the tangential velocity change is in the usual direction shown in figure 5. ‘Tho axial momentum foree then is Pam fnVoas(Voa.s—Voadb dg ® and the tangential momentum fores is Fae [Vases Wea db tg ® Sineo momentum values in the wako can be obtained most easily as differences between the wake values and the down- stream value outside the wake, it is convenient to rewrite equations (2) and (3) Vas Paaa—Faadbat f Ve Fe, 1Weas-Vaaadbdg (4) Foren VesWos—Wescdb0+ f piVeas(lFeaa—Wia,db dg @) Howover, the wake momentom force, as ealeulated from wake surveys, is = fav Pr Uf, now, the flow direction in the wake can be assumed to be the same as the average downstream flow direction, the ‘wake foreo can bo resolved into components in the axial and tangential directions. Using the same sign convention as before (Ws = Wb dg O} cos m= f nVeasVoss—Vaasdbdg ® ‘These are the integral terms in equations (4) and (5). Sub- stituting equation (7) in equation (4) and equation (8) in ‘equation (6) yields the axial and tangential force components, m2 1s follows: Pet Para= (D221) 69+ Vea Vase Vea)bg~ Fe 008 fa Fea Fare= 0.0 Wea Weanbgt Fe sin By For convenience coefficients based on 9; are used and teeyey Fag a} [AP Ves Vane~Ved |, Se seep Pe 0 Be oF 1 Va Taibo or, trata If is used to denote the angle between the resultant foreo and the tangential direction ‘Tho lift coofficiont es, and drag coofficiont e, ara the components of cr, perpendicular and parallel, respectively, to tho veotor mean velocity Wa, where Wn is the vector average of the velocities far upstream and far downstream. ‘The upstream velocity can be easily measured. ‘Tho velocity. far downstream is obtained by proper averaging of the velocities just behind the blades. Sinco the axial area con- trols the axial velocity, conservation of mass determines the axial component of the velocity far downstream. Inasmuch as there are no physical boundaries in the tangential direction to support pressure gradients, conservation of momentum controls the tangential component far downstream. ‘The discussion up to this point applies to compressible as woll as incompressible flow. For compressible flow the effect of wake mixing on pres- sures and densities makes accurate determination of the axial velocity far downstream rather tedious. In tho i compressible, two-dimensional case the downstream axial component is Vea, and the downstream tangontial com- ponent is the momentum-weighted everage of Wx ‘This tangential component can be obtained by adding to the tangential momentum of the discharge free stream tho integrated tangential momentum of the wake. ‘The inte grated tangential momentum of the wake can be determined from the tangential component of the wake coefficient. Having the correct velocity far downstream, the vector mean-velocity direction Wx can bo easily obtained. ‘Tho direction of Wm should bo determined accurately since ¢,y is very nearly perpendicular to Wq, and the value of the drag component ¢,; is sensitive to small changes in the direction of Wa:APPENDIX B CARPET-PLOTTING TECHNIQUE Sinco the earpet-plotting technique is not too well known, ‘a description of this technique will be helpful. examine first figure 114 (a). Tt will be noted thet this figure is ‘composed of five similar and separate plots. Each of these plots shows the variation of the turning sngle @ with tho angle of attack a and the inlet-air angle é for a given solidity a and a camber cx, of 0.40. Tt might be here pointed out that the tests were made at four inlet-air angles, 30°, 45°, 60°, and 70°, The leftmost plot, which ropresonts a solidity of 1.50, is constructed by plotting turning angle @ as ordinate ‘against angle of attack a as abscisse for f,=30°. Then, for a A, of 45°, the a; scale is shifted to the right a number of grid units proportional to the 15° increment in ® and the turning angles are plotted as before. ‘This procedure of shifting the a scale is followed until the range of &; values for which test deta are available hes been completed. Curves of constant angle of attack may then bo drown between the several curves of a against 6,50 thet the removal of the a abscissa seales is possible. At this point, curves of a against @ may be filled in at 5° intervals of & by using tho proper abscisen increment, ‘The plot thus constructed is called 0 GaisB: carpet. A b,au6, carpet is constructed for the next solidity of 1.25 by shifting the angle-of-attack scales to the right a number of grid units proportional to the solidity increment of 0.25 and sufficient to keep any overlapping of the 8a,2 carpets to a minimum. ‘This procedure previously described in constructing the first @a,8: carpet is then repeated. ‘The full range of solidities for which test data aro available (0.60 to 1.50) may be presented by spacing and constructing the 0,8: eaxpets on this plot called a 6,au,8,¢ carpet plot. Similar G,a,81,¢ carpet plots are then made for each of the other cambers, namely, ¢1o=0.8, 1.2, 1.5, and 1.8 shown in figures 114 (b), 114 (@), 114 (@), and 114 (0), respeetively. ‘For intermediate camber conditions, linear interpolations between G,a;,,0 carpets could be used or these #,c,Bi¢ plots could be combined into # single carpet plot to make possible a single graphical interpolation. For example, figure 114 (a) may be combined with figure 114 (b) by again shifting the angle-of-attack scales to the right a number of grid units proportional to the camber increment of 0.40. Overlapping of the 04,61, carpets can be avoided by shift- ing also the 6 ordinate seale vertically a number of grid units proportional to the camber increment of 0.49. ‘This com- bination of a vertical and a horizontal shift is facilitated by ‘tho uso of register points labeled “AB” on both plots 114 (a) and 114 (b). ‘The AB register points ean be superimposed and the grids alined. In like manner, if all of the register points are used, figures 114 (a), 114 (b), 114 (c), 114), and 114 (¢) may be assembled into a single carpet plot. Figure 115 was made by combining plots 114 (a), 114 (¢), and 114 (©) representing cambers of 04, 1.2, and 1.8, respectively. On this carpet, the design angle of attack is indicated by a dotted line and the approximate occurrence of twice mini- mum drag is indicated by a dashed line. Since the origins of tho @ sosles for cambers of 1.2 and 1.8 are shifted ver- tically a number of grid units proportional to the camber increment, the ordinate scale is no longer a truo @ scale for these higher eambers and is called Y. When an interpola~ tion is made for any camber above 0.4, the @ value may be obtained by substituting Y in the following expression: 0=Y—50(e1,.—0.4) Figures 114 (b) and 114 (@) representing eambers of 0.8 and 1.5 wore omitted from figure 118 in order to reduce the size of the plot. It-will also be noted that date are availeble for cambers of 0.8 and 1.5 at only two solidities, 1.00 and 1.50. In view of the necessity for shifting the separate plots to provide for combinations of the several variables, as ilus- trated in figure 115, the carpet plots assembled in the present bound copy are useful only as a means of demonstrating this technique. For this reason, larger separate plots for use in compressor design have been prepared and aro obtainable on request from NACA Headquarters, Washington, D. C. ‘Tho use of the carpet plots presented ean be shown best by use of an example. Generally, from a velocity-diagram calculation, the inlet-air angle, turning angle, and inlet ‘Mach number are known, and some value of solidity has been decided upon. The problem is to find the eamber ¢,., tho design angle of attack a. the pressure-riso coefficient Ap/q (one-dimensional flow being assumed), and the off- design varietion in @ with a at a constant blade setting. ‘The following design conditions are essumed: Bia55° 6° 10 85 Figuro 112 is used to locate first the intersection of the curves for 6:=55° and 24=15° on each of the four carpets representing solidities of 1.25, 1.00, 0.75, and 0.50; then a smooth curve is drawn connecting these four points which aro labeled “A,” “B,” “0,” and “D.” If the 8; and @ values hand fallen between thoso represented on the curves, these intermediate values could be located by measuring the in- crements along the abscissa, Although the design solidity - Ey12h of 1.10 falls between points A and B representing solidities of 1.25 and 1.00, respectively, points C and D for soliities of 0.75 and 0.50 are included to define more accurately the shape of the curve between points A and B. Sinco the hor- ‘zontal interval from A to B represents a solidity increment of 0.25 from 1.25 to 1.00, the point corresponding too solidity of 1.10 may be obtained by locating on the ABOD curve the point B, which has a horizontal distance from point B equal to a solidity increment of 0.10. Point E thus located indicates camber cy» of 0.87 on the ordinate scale. Next, the solidity of 1.10 and the newly found camber ex, of 0.87 are used in conjunction with figure 113 to find the design angle of attack. ‘Tho point for a camber of 0.87 is located on the e=1.10 curve between the camber of 0.8 and 0.9 curves by reading the proper horizontal increment 0f.0,07.. ‘This point indicates on the ordinate scale a design angle of attack of 10.5°. ‘The pressure-rise coefficient Ap/q, is found from figure 116 (a) by using the values M=0.65, 6=55°, and 0==15°, In figure 116 (a), the Ap/a, was caleulated on a one-dimen- sional basis, whereas in figure 116 (b) the Ap/q: was ealen- lated for a constant axial velocity. Since one-dimensional flow was assumed in this example, figure 116 (a) should be used. Employing the proper horizontal , increment, of 10° and starting at the B=45° curve, locate the Ai=55° point on the @=15° curve for each of the carpet plots rep- resenting Mach numbers of 0.60, 0.60, 0.70, and 0.80. ‘The four points thus located are designated “WF,” “G,” “H,” and “[ and are connected by a smooth curve. Next, locate on this curve point J whose horizontal distance from point @ is equal to a Mach number increment of 0.05. Point J indicates on the ordinate scale pressure-rise coefficient of 0.590. REPORT 1368—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ‘Tho last step in this sample problem is the prediction of an af curve. at a constant blade setting. ‘Tho offlesign carpet (Bg. 115) is used to predict this a,0 curve for the blade seotion having a camber oy. of 0.87, a solidity e of 1.10, and a. blade setting of 44.5°, ‘The blade sotting is the differ- ence between the inlet-air angle and the angle of attack, or 55°—10.5°=44.5° at the design condition. In figure 116, curves labeled “a” and “b” representing this constant blade sotting of 44.5° are drawn on the éai,8: plots for ex. of 0.4 at the solidities of 1.25 and 1.00, respectively. Curve ¢ is then interpolated for the solidity of 1.10 by the use of the correct solidity ineremont along the abscissa. As can bbe seen in the example, this interpolation is aided by drawing Ddotwweon curves « and b lines of constant angle of attack at values of 6°, 8°, 10°, 12°, and 14°. A similar interpolation is then accomplished for a camber of 1.2, which produces curves d, e, and f. A linear interpolation for the intermedi- ate camber of 0.87 is mado between curves ¢ and f to obtain curve g, which shows tho variation of ¥ with a for the design camber and solidity. ‘The ¥ values may be converted to @ values by using the relationship ¥—50(¢,.-0.4) It has been found that linear interpolations between any ‘two cambers of figure 116 produce design turning angles which agreo with the design carpot plot within 1.0°. If greater accuracy is desired, a faired curve between tho threo cambers should be used. In figuro 116, tho design angle of attack is indicated by a short-dashed line and the approximate occurrence of twice minimum drag is indicated by a long-dashed line. REFERENCES 1. Bria, Jobn R., and Emery, James C.: Effect of Tunnel Contigura- on and Testing Technique on Cascade Performance. NACA Rep. 1016, 1951. Supersedes NACA TN 2028) 2, Westphal, Willard R., and Godwin, Wiliam TR Comparison of NACA’ 65-Serias Compressor-Blade Pressure Disteibutions and Performance ina Rotor and in Cascade, NAGA TN 3808, 1958. (upersedes NACA RAC 51820) 8, Bogdonoff, Seymour M., and Bogdonol, Hacriet B.: Blade Design ‘Data fot Asial-Flow Fans and Compressors, NACA WR 1-635, 1045, (Formerly NACA ACR L5¥07a) 4, Bogdonoff, Seymour AL, and Hess, Eugene E.: Asial-Flow Fan ‘and Compressor Blade Design Data at 62.5" Stogger and Further ‘Verifeation of Cascade Data by Rotor Tests. NACA 'TN 1271, oi, 5. Abbott, Im HL, von Doenhol, Albert ‘Suminary of Alrfll Data. "NACA Rep. 824, 1045, NACA WR 1-560), ©. Yates, A Hs ‘Cumpets’ and ‘Lattices’ vol. XVIII, no. 208, Jun, 1946, pp. 8-9. 7, Bursoall, Wiliam J., and Loftin, Laureneo 1, Jr: Exporimontal Investigation of Localized Regions of Laminia-Boundary-Layer Separation. NACA TN 2898, 1051, 8, Howell, A. R: Design of Axial Compressors. Lectures on the Development of the British Gas Turbino Jet Unit Published in War Bmorgeney Issue No. 12 of tho Institution of Mochanioal Engineors. ASME, Reprint, Jun. 647, pp. 452-462. and Stivers, Louis 8, Jn (Guportedes Aircraft. Bnglneoring,SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SHRIUS COMPRESSOR BLADES AT LOW SPEEDS Epo. ' Baio corel se Seting chamber Frovae 1- -Vertial cross section of to-dimensional low-epeed casoade tunnel. Fes arora. j eae \ Fravas 2—Photograph of Langley Sinch cascade tunnel with ‘portions of one side removed to show porous surfaces, i Foarsl 3 C ° a a 12 1g 20 2a 28 Presid, bq Fraums 3.—Porosity characterstion of the ealendered mone iter loth used In this investigntion. 725tea 65-150 roe oe nach 65-1010 Short Qos p30 i : T “Sons Gone oe nica 65-210 org ert Stent “me ica 65-810 Nac 65 (2010 vga s Toop @) NAGA 65~ (12110 (o) NACA 65- (2710 “Tangent (@) Lower cambered sections. Angle between chord line and tangent to lower surface as shown for tho various soetions. Frovaz 4.—Blade teotions tasted in this investigation. () Higher cambered sections, Chord line and tangent lino soineldont, Figure 4—Coneluded. ‘Fiune 5—Typieal vector diagram for a compressor rotor.SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIES COMPRESSOR BLADES AT LOW SPEEDS 727 2 Comat sree a 2 Geren race 3 Cina oan a 4 s roy (0) ay = 30%, 02.4" 5 { - ma eo 2a oe | fm a-30, 0-8 eal u s Ss s pao (d) a+ 8.0%; 8+6.9° a en 7an, 8-99 Bsn aq ACen 49 aa ql ; J s . t 1d . @ gethr, 6-934 0% 40-25 85 100 in) ays170r; @=143° 02 a0 65 60 100 ceed or sy 40.307 wt | 20 {06 | 2 +05 ft 2 10, 40445 Bde Fy 44 ow | 0° doses o of J-t0 -Jo2 “t 1-20 or i O48 6 20 oy oa 26 ( Section characteristios; flagged symbol indicates Teading-edge roughness Fravns 6.—Dladesurfaco pressure distributions and blade section characteristics for the easeade combination, 6:=30°, 7=1.00, and blade seetion, NACA 65-010. © ayr150y Or16.6" 06 ao 20 80 100 a i) aj-210%; 02210" 080 a0 65 80 05 cat crt ep 7 60 05 vat ol Iso os al si 40. fos ‘ba 30, 402 ey aes Fe + abs 120” 01 ey of 2 w Jo al ° 401 4} ne oa 8 ee 20 09 @) Scotion characteristics; laggod aymbol indicates Tending edge roughness. Fiavns 7.—Blade-surfaco pressure distributions and blade ‘seotion charactriaties for the eascade combination, Bi=30", (721.00, and biade seotion, NACA 05~10.REPORT 136§—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 3 Genel aries T ap ee | 2 Corer arTece a ea} | [eee ° ve = le } 2a a 28 ‘88° . 2a s aero i al ry " . | h. @ aah ee ana7 Geter @ sere t 2 @ alin Oe2ur Fy 40 2 sq a : KJ 1 2a Ly ehred “p P e187; 8-205" 0 ej2217%; 64280" TN 0-85 40-65 8 G9 -a a a0 80 16S ‘| Pacer See = 20-08 eae 20, O32 feta q ot a aoe oa OD BO WD 2a ro. for oF ele alee i eo 507 ats wo Sau Lo fees Fray sel sls ay leo _ . so c cog bey q aa eo ot | 0 J E jzo. oe ae so hoe ab oa Lee 10 Fou ae a e leo “Jor ot ol oto = wl 4 jo to So ae 1S Ie 20 a 04 8 ie 16 2 2 os oe dey onsen (@ Section characteristics; lagged aymbol indicates leading-edge roughness. Provan 8—Disdesurface pressure distributions and blade ‘section characteristics for the eagcade combination, 7m 1.00, and blade section, NACA 65-810, A= 80%, Fiauns 0.—Bindesurfaco prossure distributions and blade tion charactersties for the easeade combination, f= 30°, (@ Seotion cheraotorster; lagged ayinbol indicates Tening-edge roughness 1.00, and blado section, NACA 65~(12)10.SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIES COMPRESSOR BLADES AT Low SPEEDS 729 3 Gina —— ea “ae | Sena 1.6 — tot fod e f s }— 3y—F ‘| tS @ @ ayri6.0% 06297 d fe) 0-50", O15 a rihor, 8-232" 2a al ve 1 + s fA : is ‘ (0) a, "81"; B= We arly 02250" (0) y+ 12.0%; al aa ue 16 Ps . ‘ cee i a i s | 1 ‘ in | Osan beer | nas ewer fo) arty 0-008 | [fin «j-38a0, Braco eee E ore age | fee ee ree Ee, to 245 i ro or aal 9f 24 70 Jos Gy f on die doe Be alse ie ad Sa Sad a po ee - 36] 7 eat jong sd 2 fetes sat eo” os A low aE fo ee a ie 16} Al 10 or a 4) zo 08 pie le Ne 2 3 ee 16 20 24 28 Be 36° _ _ 3S 1 20 et es 8 36° ° (@) Seotion characteristics; flagged symbol indicates: (g) Seotion characteristics; flagged symbol indicates os ster Figunn 18—Blado-surface pressure distributions and blade ‘seotion charactoritics for tho cascade combination, 6,30", ‘7=11.60, and blade evotion, NACA 85-(19)10. Fiounz 19.—Bisdo-eurface pressure distributions and blade ‘seotion characteristics for the cascade combination, f;=30°, ‘71.50. and blade section, NACA 65-(15)10.734, ‘REPORT 1368—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 3 Sier action 2 Gina atten eal 8 Gene tatote (@19)= 00%, 6+ 05" | laa ooee aa 9 ad 2 : Pee ad rk ea s \ jo son, b65| | floss, boos oa ae oe ao eg 8 ae Bo Becet Ser FY 5 or d {e) 0*25.0%; 0+ 41.6" Uf) @y# 31.0%; 847.3" ae ia oe 85-05 80 tot ae a Panes ec d os ar 10 co 306 a aat shel loa so 405 4 |°* on ee 8400 Fay 4 ons wok ss so, ote A los % asst a, 44 es sot so os 2 ob le st 20 doe 28] 5] 10 01 ‘xe kee 0 | a ge 20 ee se? Se a + e 2, 6° ° ons ane (e Seton charset; tagged symbol Insts () Seolion caracteratn tagged pmb ndaten Icadingedge rupisees icadingodgs ropes, ‘Frovan 20.—Blade-curface preasure distributions and blade ‘section cheracterlaties for the eassade combination, pi=20%, {50, and blade section, NACA 65-(18)10 Fravaz 21.—Blade-aurfaco pressure distributions and blade ‘section characteistieg for the easeade combination, B.=46%, 7=0.50, and blade section, NACA 65-10,SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 6§-SERINS COMPRESSOR BLADES AT LOW SPEEDS 735, eae ae 2.4} 2 a \ s rear; Bae zal s 4 L 49 Fe al s me u othe f AL fo ss ota | Pita, 9a 0 AOC 80 0a a0 aD Poca ch sr 13 90 08 es wt ule LA leo jor ak if ot Lo os a leo. fos Sy 840g 20} 4} {504-404 ent % rc lio os 8 fso. foe q bo {ou a © Mae ee a? eyes @) Section characteristics. Fiounn 22—Biadewurface pressure distributions and blade seetion characteristics for tho eascade combination, £1=45%, ‘#=0.50, and blade section, NACA 65-(12)10 2 Care ats 2a a 108" Re 7 ! (oar for ayer Gaia “ aa ahaa! s oe ‘thay 2U7%, 0-184 ‘0 20 40 65 80 105 020-405-080 10, Pacer et Br may 708 x ee Heo or oa 2a 70 os a 0 os Bde Fay, Fy ont 20 20° losers a re 10} 20 Jos i 30 oe 4 20 or al wo to ee eee 809 @ Section cheractorstes; Aagged symbol indicates leading-edge roughness, Fravas 28.—Bladesurfaco pressure disteibutions and blade seotion characteristics for the eascade combination, #15, ‘©0.50, and blado exetion, NACA 85-(18)10,736 REPORT 1808—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 2 Ciel aati 2al @ 324 oh A CEES a fed ayn 14.0%; 8-12. i) 922005 8 15.8" 08 4086 800 “0-30 40-€5 80105 Porn 2a 50-07 2a 40-408 2] 30 }08 '6| 2 40% e ea dy on re 10 Jos at 0 +02 | 0 or a ° {@) Section characteristics; Sagged symbol indicates leading-edge roughness. Fiouan 24—Bladesurface pressure distributions and blade ‘eotion characteristics for the easeade combination, 845°, 0.75, and blade weetion, NACA 65-S10. scare 24 2 td |_| : rN area Oca ea sshd 7 E 4 i @ Te “ ad 2a ° he d a von 8-22) 085 5-05 a a0 20 5-66 8 HO Facet Sad eeu 20 03 al wo 0 jor at sl co os at co os, Bees bay 44 ih "7 140 00% wt al 20. Jos a | 20 oe <4 10 Jor a ee ea ee ° or) ]) Seotion charactoristiea; Nagged symbol indicates leading-edge roughness. Frouns 25.—Blade-surface pressure distributions and blade ‘ection characteristics for the eateade combination, p=45°, ‘760.76, and blade avction, NACA 65-(1)10.SYSYEMATIO TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIES COMPRESSOR BLADES AT LOW SPEEDS 737 3 Ginet aries / eal asa 3 imal aris a a rie 6 [| c reer , 7 | s ch t os gf 30) OIF 4 tera 2a @ acto; Ota | [eo arto 6203 24 7 s bo Oa fis i sy s © s60% 3 a o| aq BH ayrl4or, 62235 | fl oye l60%, 82250" q 40] 32) aa ea [ ; k 2 PTF 3 } : pon [apes fr asisor, boar 210%; 89286 pie palepcogaet deg (@ Seotion charastorstes; Aagged symbol indleates Teading-edge roughness Frouns 28—Blade-aurfaco pressure distributions and blade ‘rotion characteristics for tho eascado combination, 6)=45%, (70.75, and blade eestion, NACA 65-(18)10. P 1 ab ad os, awa bey + ot” a wt at | 0 on los"! 0 2 4065 80050 -® a0 60 60 100 reel ord = os os os Bae ae 20 dor sere ao Jo a deg @) Seotion characteristics. Frouns 27.—Blede-surfaco pressure distsibutions and blade ‘section ebaraotoristies for the easeado combination, f,=45°, (00, snd blade seation, NACA 65-010.738 REPORT 1368—NATIONAL ADVISORY COMMITTEE FOR ARRONAUTICS 2 Gia ice zal 2 Sey seo 2 Cine ice \ pl 8 Sonar Safes a |_| a 57, 8-107" s erepel (0) ay 575 Pres ao oy 50%, OTH hal s pes 8 [" oy 87%, Bie? (i gy87%, 0-180" I 4 @ ro set 4 JE 24 32 s vl 2. t I i s 4 : 187% 8 209" fo ia @ ayn l40, + ase 0% 40 eo 86 160 KO 2,195"; O« 173" 2% a0 60 80 160 od 50 705 z ‘6 % lao os rd so ose 869 $4 a q 0” foot 4 1 don ove a : Cr a “0 () Section charsoterstca, Fioumn 28—Biade-surface pressure distributions and blade saotion characteriaties for the eazeado combination, 09, and blade section, NAGA 65-410. 0 20-0 65-80 09 080 40 60 66 100 Parent cord sect xe e008 ze ot 9) 324] lo dor 8 a ol iso fo8 nie at 7 lo fos gece | ey fat if 5 hod 404 ane a fo fos o 20 doe 4 10 fox @ Section charaotoriatos, Fieuns 29—Dladesurface pressure! distributions and. blade section cheractoristios for the cazoado combination, y= 40%, 1.00, and blade seetion, NACA 65-810.SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIBS COMPRESSOR BLADES AT LOW SPEEDS 739 3 Came suis 2a} oon jo) a= 70%, 8 165° 70) j= 10", 84208" 2. s 4 4 peitiees, we] | flaw sjror, 0-20 a 4 +1 oie ad 2 : x a L od ttt tt | ea s bt aa ihe ae Faw oO a rae arto oo Toc ed x po 08 sl 70 or al co fos a fo os 4,409 16} [40 4 osand | 30 jos at leo “joe 4| to for al oto () Seotion characteristics. Fiauae 30—Blade-surfaco pressure distributions and blade tection characteristics for the eascade carabination, 7211.00, and blade section, NACA 65-(12)10. 48°, lo 9,# 210%, @+ 300" 0 20 0-60 80 100 aj 270", 8» 324 25 40 66 80 160 ef 32 dea (@ Section characteristics; lagged symbot indicates leading-edge roughness. Figoxs 31—Blade-surface pressure distributions and blade ‘ootion characteristics for tho cascade combination, 6, 731.00, and blade seotion, NACA 65-(15)10.REPORT 1868—NATIONAL ADVISORY COMMITEE FOR ABRONAUTICS 3 Gael are 2a 3 Ee ae 1g 3 2al wg ai s a we arity BS | fio ania, OTe 4q 2a 5 PS I u 2 | [Fibe-250y oe of a 40-608 Wo 8 ade 85 Sos wor 0 407 af leo Jo xf io) so Jos al | 40, lose, Bees bo cae aah a 30” fos ook | 20 oe wh 10 Jor 7 yi w See ee a5 ae za? aya (@) Seotion cherasteriaticn, Fiounn 32—Blade-surfaco pressure distributions and blade stotion ebaructeratcs for the cascade combination, #1=45°, 0, und blade cection, NACA 05-(18)10. 2al Beacorss oT 2a (_ayri95%) 6233.0" 7 32 2a sey E aef 13 eo ab 10 aof uy so sf 10) so a0 bay seb af a0 al ol 30 al zo ar si 0 a Hie ‘See 20 ee ee 8? oc) @ Section charactoristies. ‘Fiouar 33,—Blade-surfaco pressure dlstributions and blade ‘eotion charactaristice for the eageade combination, f= 45%, ‘21.00, and blade sootion, NACA 66-(21)10.SYSTEMATIC TWO-DIAGINSIONAL CASCADE '™HSTS OF NACA 65-SERIBS COMPRESSOR BLADES AT LOW SPEEDS 741 (2 Cima eatice eal 8 Cone aefose 304° | fa) a)-2007 Poss" 2a lod ay +240", 8-38 gy 2. 2h 2a at lo Joe sok lo. Jos sek id 0 ote Odea F Sy 34 os set lao” oseay at lo oe ad | eo ox pel 2 See 0a as Bea ( Seotion characteristics; lxgged symbol indleates leading-edge roughness. Frounn 34.—Bindesurfaco pressure distributions and blade seotion characteristic for tho cascade sombination, = 45°, 1.00, and blade seotion, NACA 05-(24)10. Ty 0516 = Gone oatoce [Pde veh nH ErPRHH Pet recatae, | [a= 270%; 0-aaee, ya 210% B37 49] (@ay=280"; 82443, 32 2 a Poa f Ps, q 0 20-40-20 BOT O20 a0 6080 Te hor cor 08 ssf | lor sf Los af ul los ea aaa fay nt a Jos «of wo los ie af | Loe x ol tor fal Tse a a a ade © Section characteristic; flagged eymbol indieatea leading-edge roughness. Frovrs 35~Bladewurface pressure distributions and blade swotion characteristics for the cascade combi 71.00, and blade seation, NACA 65-(27)10,2al Sey spore 6 si | Oh Lor Warr or aa al \ s Ptetel A Jemtets @ a5; O15" ase a 3a al s EBS iia (@ ay=170% O° 17.5" fey ay=220%; =20.7° 0 ao wo 0 WO 0-H 40 G80 OO chet ae Ie 70497 zt slog 60 108 8 oa aot slat 50 405 we 4 40. 4.04 ey Bdeg Fay 44 ow ret” 3 30° fos a 2] 20 foe aoa 10 for eo wwe 7 0 (@) Section characteristics; Hagged symbol indicates Tending-edge roughness. Piovur 30—Bladesurface pressure distributions and blade ‘ection characteristics for tho cascade combination, Bi=45%, 751.26, and blade seotion, NACA 65-A10. 2 Chm aattce a 8 Senco 1s} s 7 | _¥ fo) aj 7s 6215.1 | [a oi12.6%; 8-209" 2a ' = a @ oir 18.1%; 8-26. “H ‘ Gar ery fe) a 240; B= 3h Ana 200 0 ao eo 80 020 a0 60-8 10D Pecan ed Br OFT 80 308 | i 70 or 2a Jeo jos aal £0 os 34002 8d oF £ 4034.04 ond % '6| 20 Jos 12] 20 oe a to Jor al ° @ Section characteristics; Mugged symbol indicates leading-edge roughness. Ficues 37 —Bladesurface pressure distributions and blade fection characleratics for the cascade combination, fi=45°, 25, and blade seation, NACA 65-(12)10,SYSTEMATIC TWO-DIMENSIONAL CASCADE TRSTS OF NACA 65-SHRIES COMPRESSOR BLADES AT LOW SPEEDS 743 3 Set sae TI 2a a tee (hay + OR, 06 235 d Fe eal FS :| 5 (@) 90.0%; O=-1.5" Os, 2at 1s} s pot @ 4250% 664 es 32} 2a 2 Cine erica 8 Goreme erace 2a Prorss ey _ a co oe “tad bo Jor a Io oe ae + me ef | at 4 loos at eo oe oot lo do APE eo 20 Be BS BE 36? ay oe9 (@ Section charaoteristis; faggod symbol sndicates leading-edge roughness. Frovne 38.—Blade-surface presture distsibutions and blade ‘seotion characteristics for the easrade combination, #,=45°, 71.35, and blade seetion, NACA 65-(18)10. ‘2007-00-40 150 06 16} \40. -105 I30, {04 eq: 5 one 20% 103444 to oe = o Jor a a ay 9 ( Section characteristics; Augged symbol indicates leading-edge roughnese. Frovs 30.—Biade-surfeco pressure distributions and blade tion ebaracteritios for the easeade combination, Bi=15°, t7=1.50, and blade section, NACA 85-010.744 REPORT 1363—NATIONAL ADVISORY COMMITTEE FOR ABRONAUPICS g Comer srtoce TT 2.4] [eee 2.4) o at op | dh Gl 4 ; ms : J 5 . ett A 5 7 4 , sd C a 2 7 « K : i" | Pe] 4 0 ao 6 66 G0 a0 60 0 OO cent ap sop 60 307 ba woe a5 20 06 oa 6 4 40 os oF , x 20 , oa 65 745% 20" os ces 10 foe of 0 Jo £ ai 7 ete amepicuisnciets @ a dog (@) Section charncteristcs; Aagged symbol indicates leading-edge roughness. Fiouas 40—Biadeaurface pressure distributions and blade ‘section characteristics for the cascade combination, #=45°, 7-11.60, and blade seotion, NACA 65-410. fy a= 257°; 8-206 0 85 a0 65 8065 ent © 0a 40-€o- 6016 Perot 70307 Js0 Jos js0 os 140 , Joa % Fy 307 os Gh 20 00 10 Jor ‘ a al ° So 4 e 12 16 2 24 28° a dog © Section characteristics; Haggod symbol indleates leading-edge roughness. Fioume 41—Bladesurfacs preasure dlteibutions and blade sootion characterietios for the enseade corabination, £.=46%, ‘71.50, end blade section, NACA 05-810,SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIES COMPRESSOR BLADES AT LOW SPEEDS 745 ea 24 = i Te] ‘ ree _ : , : x : * : ; 5 cS (e) area, 830.6" 0 2 40 eo 80 0 36 60 06 xf a so los aa] {40 104 deg F ou i zat 5} 130 $1.08 ond % aot al lo oo tof a} 10 Jor i a a a ausea @ Section characteristics; flagged symbol indiestes leading-edge roughness. Fiavan 42—Biadeurfaco pressure distributions and blade section obaracteristics for tho cascade combination, #,=45°, 1,50, and blade seation, NACA 65-(12)10, Reacerad | “arte a i 2 4 : a ‘ : ler a; = 280", 0 5" on 0 eo 8 wy ft) a =320", 8-409" 25 40 60 6060 ect Sot apo 7 or aot} Sait {60 los st ola os sf | Loy 8,603 +] a zak loses at loz aot lor 's [2a a a 0 ones (© Section charactvntin; Gagged aymbot indeses Tesdingedge roughness. Provan 43.—Blede-surface pressure distributions and blade seotion charactoristics for the easeade combination, #:=45°, 2-150, and blade seotion, NACA 65-1510.746 REPORT 1868—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Toate (© Comex sa foce 2a 2.4] *. 4 g ve s 7 uh ‘| Stele ‘ (b) y+ 20.0%; 8 +36.8° (0) 714.0%; 8=28.7° (e) a) #19.0°; = 33.8% i d 2a ; aa J \ . : f . : q SN propo pak @ e240; Baur | [la «ynz70%, Oo Fa,r a a a a ea t 2a {| she s i LY J \ - i 3803 8-51 crane | [Paasseor teat oa ey Boe Ho wo 80 of ab a5 eh oo Robo Ho 0 BS 1D seu 20 08 Feces a ar 0, sb 20 dor af ol af ol eo os af at 20 loses 8de9 bay $4 4a xf, | “of 20° loo aon bay 3ef 6} ach | 30 Jo at 4 wt Jo oe at al at a 10 Jor E i @ H a i a ee ee Age 20 a 2s 82 36 40° ~ ° a outed (© Section cbaroteraticn; Sagged. symbol indlntos on dng- (@ Sesion shurctrties Aagged symbol incense igo rooghne sla eymtolneatr high Reyna umber Traine ede roughness Flours 44-—Bladesurface preanre distributions and blade Yiovan 45—Sladecurace preenro diebutlons and wade eatin charter fort cuca sombintion, A=" ‘euion thnacforts for fee easade combination, A=, Soo, and blade section, NAGA 68 0)10, S10" and Bade section, NACA ao-()10,SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SBRIES COMPRESSOR BLADES AT LOW SPEEDS 747 esccas 2al (©), 2/9160 122.0% 5 da 2904 lar ad a ea > FELEET REC vet Cl © | Piacaorecins Oe aes a a —— oa a6 wo 8S Recon Sad co 190 310 a led, 20 Loe cok alec 20 oe see 10 Jor st 19 60 , jose de as ond Bor so? dost wk a 10 Jos wok 7 x0 Jos xt 20 oe xb 10 doi th 2 Age 20-26 WB Be 36 a0 4a? ~ O oc (@ Seotion charactoristics; Aogged symbol indicates leading-edge roughness, Frouns 40,—Dlade-curface preasure distributions and blade ‘eotion characteristics for the eascede combination, fim 48°, 50, and blade section, NACA 65-(24)10. Seat | Eee us| 3 a 02% @ ea \ _ 5 d @ ar4ey 4d a wu : it Oana oe | Parity Boaz 0 & a5 60-80 080 ad 60 80 WO cent 2p a 60 4.08 bleep zo zfoay i007 [ “ec iL of so 106 eb | 50 | -195 aaa Bee bay oy on ak al Jao. 104 % ab | 130 103. of | j20 {02 ab '0 Jor oto Oman ere eu6 oes () Section charaoterstos; flagged aymbol indicates leading-edge roughness. Fiauas 47—Bladesurfaco pressure distributions and blade section characteristic for tho earoade combination, fi=60", 0.80, and blade esotion, NACA 65-410.748 REPORT 1865—NATIONAL ADVISORY COMMITTEE FOR ABRONADTICS ae, 5 moe 24 oy ea 3 Gace an pj || u ; wh Ne s s 4 53 to) aor 8 24 =~ ve s 1 1A 4 4 aa — sq t 2d s 9| GH ays7 Hersey Brie oF 8 aoe Bea “93 40 60 80 Ko ge eae ame eee ee ee i 0406 mr 12 eo 07 2 i ae ou ro os zhu Lal los 16+ 19] joo 105 a sf io 4 *ees 12] [50-104 Gay aden Pay iy + of eden FO 3} ond rf veut 4 lao” os = of l2o oe a4 | Iso Joe 28 + ' lao or ab 2% Le Jor al “TTI, 15 rad a | ° a one audey ©) Section charactristion;Sagged symbol indicatea (©) Seotion characterises; lagged symbol indicates Teadng-edge roughnen, Tending-elge roughness, Miaune 48.—Dladesurfece preasuro distributions and blade soction characteritis forthe eascade combination, f.=60°, 2=0.80, and blado section, NACA 68-(12)10. ‘Freons 49—Bladesurfaoo pressure distribution and blade ‘ection charaotoristis for the caseade combination, Bi=O0%, 0.50, and blade section, NACA 65-(18)10,SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIES COMPRESSOR BLADES AT LOW sPHEDS 749. 2 Cana aatice 8 Sone a 33; OF40 (@) a= 93 fg = 9.8%; O24 0 20-40-60 80 105 Pe 9 ay #18.1°; 8 = 110° 020 a0 6 80 160 ct ced mr ae 70 307 oat wef} 23 eo Jos 12) cs 50 4.05 at a 20403 ey aden bey Hom abs 20" os of 2 20 oe af a 10 dor [ io) le.¢| 7 ° a 6 1@ 16 a? 2 ay ag @ Scotion charactoristiea; Sagged symbol indicates leading-edge roughness. Frouns 60.—Dlade-surfaco premuro distributions and blade rection characteristics for the cageade combination, $1=60°, ‘em0.76, and blade sootion, NACA 65-410. TE a ro Gq Toy ad x 5 5 iol HN Lebodols INL lo wrt32; 62 162° 0 26-40-6080 165 0 20 45-60 80 160 ord aa 1 170 406 of | Jeo Jos if | 50" oa eigen 6, deg 12-4 7] }40 $4103 ond % sts 30 {02 20 at | Ts lot 4 8 2 1 a a? ay dog (@ Section characteristica; aged symbol indicates leading-edge roughness. ‘Frouwx 51—Bladesurface pressure distributions and blade stetion charactorstin for the oaseade combination, pi=60%, 0.75, and blade section, NACA 85-(12)10,750 REPORT 1808—NATIONAL ADVISORY COMBITTEB FOR AERONAUTICS sonar al sai ce we ; s ‘el aa s| s Ht o ze A A zal ose | [owes ise 24] ‘| 5 . 4 ho s @ aac oor | [@ a a9] fo par ber | Laser 40) Ls 32] + a s _| sf 2a + s L ' q © ar60y 05a | fin acter 8-10 08a 65 85 WO 0-25 90 BBS 1S Sed ar 5 20-08 se E "te Pe eR, fw ao 6 80 0 wt af oa 0 or Paley Bp ro 498 we ale t 30 106 zak to} loo {os 20 405 ey Hos zat 9 150 oy 10" 404% don 4 od Eb ok a 04 os Ss to © os 16} js -j02 on le lo Jor ed oo ich = jo Jo : al Se eae ae 8 Spee ae HOO cen a) @ Section characteristics; flagged symbol indicates (@ ‘Seaton shart Teadingigo roughness. Frouns 62—Blado-curfae pressure distributions and blade ‘eotion cbaraoteriatia for the caacade combination, 6.60%, 70.75, and blade section, NACA 65-(28)10. Fravar 53.—Blade-surfaco pressure distributions and blade ‘teotion characteristi for the cascade combination, = 60", .00; and blade section, NACA 05-010.SYSTEMATIC TWO-DIRIENSIONAL CASCADE TESYS OX/NACA G5-SBRIES COMPRESSOR BLADES Am Low avzmDs 751 3 inate ea) [eee \ s a| eh splos Or Wayeaor, Oras 4 2a 1¢ s ai @ ayn 60%, 0-68" 4 32] 2 \ me @ a 105%; 8100" | fin op 150", Brie 0 40-60 80100, 20 406080 160 Pera eho arom eo 506 aol | 20 oo hs so dose,, aos bay Hoos ab 20° os et | 20 Joe foe © for ee Lla Ow fe ooo ( Section chasncteristoe. Frouns 54—Dlade-surface prewure distributions and Blade seotion cheractaristica for the eagsade combination, p:—00% 71,00, and blade section, NACA 85-410. 2080700 —80 2 Cine eatin ea 8 Senco sfc ss bpeey a (ae67; 8102" 115 2al (Or (0 ail 4a} Fe Jorasta7 esti] _fisrmciermy aeibo oe oy 85 bo 08 40-80 bo Pecet Se ars 60408 zh 7 50 os sts 0 403 4, aa Fey $4 ont 1 5] 130 402 4 20 for ee a Agee ead? ay. do () Seotion charactoristics, Frovar 55—Dladesurface pressure disteibutions and blade section characteristis for the cascade combination, #60", 00, and blade eeetion, NAGA 65-810,752 REPORT 1808—NATIONAL ADVISORY COMMITTEE FOR ABRONAUTICS at arti | 8 Gonaare surface 16 i bot s (0) 074.2%; 0°90" 2178.27; 0135" | 2a ‘ = s @ @y=120 32| aa s Mf) ay 217%, 9922.5" 0 & 46-60 80 Oo & a0 6 a0 OO Parent chord Rr to 70 3.06 zat | co 405 ast ol 50 4.04 Gay Odeg Fu od ak 7 J4o5 fos if 6} so 02 iat 5 20 01 a) ise a a a a cu deg @ Section characterstin; Aagged syrabol indicates leading-edge roughness; solid symbol indleates high Reynolds number. Fiounn 50.—Blade-surfaco pieagure distributions and blade ‘ection obaracteristies for the cascade combination, #.=60°, 00, and blado section, NACA 65-(12)10. (©) a= 80% 0 189° 2a 3 See ace rT . ra 2 ee A po dod I T] 5 2a fe) a, mor; Be2ise | [ta tear Ozer . a ms Pe 4 joa, ies o-eo | [i a Be 32| 2a s A 40 60 80 0 20 40 60 60 100 Sud 2s 60 308, 8 J eof 288 bo foe = : zo 3 lose, ada be £4 ons to | Blo loo% x 0 ou t .# ew Caamcieawiggoe gc 189 @ Section characteristic; lagged symbol indicates leading-edge roughness. Ficune 57-—Bladesurtace preasure distributions and blado ‘seetion charateristies for the cascade combination, 6-60", ‘71.00, and blade section, NACA 65-(16)10.SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIES COMPRESSOR BLADES AT LOW SPEEDS — 753 2 Gime ari 2al aye)40%, Oe8gr 2al r © euieay Pam 4d] sak sto ote | Pi erze0 bts 0 B's co GO WO 92 40 60 80 Puce ce ze 10 60 3.08 al Ns oe ab a leo 2.003 4 2 Hip ‘6 jo joe 38 el 2 lio or a a . rr a aude9 Section characteristics; Dagged symbol indicates Teading-edge roughness. Fraunn 68—Bladesusface pressure distributions and blade ‘ection eharacterstice for tho eaccade combinstion, #00", 1.00, and blade saction, NACA 06-(18)10. wien, O10 | [Baye T60r, 8 209" 8 Ganesan 3 eaniee see ¢ s d 2a (as180% 0-266 | [Waei9sr, a-a71 i s (ne 2a" 3a 2a 025-408 85K 020 40-60-80 105 Percent ca 2p 19) 704.07 06 50.05 140 -}.04 ens 34 ans j20 403 leo 402 10 401 ee fs 2 ie ao ey, 09 @ Seotion charactoritics; lagged symbol indicates leading-edge roughness. Fraune 60—Blade-murface pressure distributions and blade ‘ection eheraoteristcs for tho easoade combination, 6,=00°, (e1.00, and blade aoetion, NACA 6&-(21)10154 REPORT 126§—NATIONAL ADVISORY COMMUTTEE FOR ABRONAUTICS seen oa Eee ad 1 2d = : s 10} 4 \ Gates RE me arorersr{ | [wanier;e a oa 20} ‘ 5 al NN a BeBe] fo acter, bBo 8 ad ao ad ad ed ’ s, a 2d fl vol ators Ore fo cao bias |_ Piraeus 15m ob DED OS Oe ww oa aes Be wo GBD wo et cd = mr 4 co 308 Bp ao —- oe y [ aa 23, og 7 ie 0) 0 150 Joa al oa po ee r [se ab | J2o. fone, wk | 0 {03 4, ada bay aye e000 bes 771 $4 od zor 6 ls 03%! Bt 2 30 Joe ie lo oe i bt 4 0 or 2 jo tox a fe ale ol ° 2 eo 12 ie? ° 8 35-42-16 2024 20" 7 a, deg, By deg (o Stoton sarctertin;Sggedeymbolindenten (e) Sesion charter; tagged bel Inieten Tredgcede: uric, caine ounce Fravns 60,—Blade-surface pressure distributions and blade ‘eetion characteristics for the cascado combination, f= 60°, (om1-35, and blade eoetion, NACA 65-410. Froone 61.—Biade-surfaco pressure distributions and blade ‘section charasteristes for the easoado combination, f,=00°, 1.25, and blade section, NACA 65-(12)10.SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIES COMPRUSSOR BLADES AT LOW SPEEDS 2 Goer aie 3 Gena oat A 5 Gt eta ea [sematee i rg s we : data ato ® ae fm 2-00 2al 2d 5 5 8 4 a @ ania, bbe |_| e200, be Ceres 49} 4 al aq 2a eal s s u aT 4 fe aconon Blow | [0 anengn dt fost, arr | fio arin; anor 0 ® 0-608 1609 #0 40 60-80 00 ee Pacer or Puce ch wr : 7 08 zor 4 po yor + od st a [28 ee it 3] > tos L tf | et to os ob AL Ee de, “ Bees bey 4] we Been orgy | 0 2 fotos 2b 6 tao 103% we [ + 0 tos aol | tEL Ali toe of = j@ 4% wh a zo or “po 20 jor | 1 J : a we J 7 20 a rr a a,c @ Scotion cheracteristcs; flagged symbol indicates Teading-edge roughness ‘Fiounr 62.—Blade-surfaco preseure distributions avd blade section charactoristics for the easeado combination, #.=60°, 71.25, and blade seotlon, NACA 65-(18)10. ( Section characterstin; Saggad aymbol indieates leading- edge roughness; wolld ‘symbol indicates high Reynolds umber, Fiavae 63.—Bladosurface pressure distributions and blade sostion characteristics for the cascade combination, 260°, o=1.80, and blade section, NACA 65-010. 755756 ‘REPORT 1308—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS e ee) | Coo ee [| ial . CI °F : x 1 ; mt “TCL et @ ae80% 5 Jos me | a2] ‘ aa e _ ey : a aa a : RCC ‘ cet aa bel | les aS en eer caret eyceg @) Section cheracteriatics. Ficumn 64.—Blado-curface pressure dlistributions and blade ‘ection characteristic for the eassade combination, #:=00°, (@=1.50, and blade seotion, NACA 65-410. 0 2 a0 85-80 BS 0 a0 6 80 WO Porat a aap 7 706 2 co os, aon jones | Ja" Jos 9 lo oe 4 lo jot ol ud ° ee 6 20 aO ay dog G@) Seotion characteristics. rao 65.—Blado-curface pressure distrbutions and blade ‘ction charaoteriatce for the ensoade combination, #100", 50, and blade section, NACA 05-810.SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIBS COMPRUSSOR BLADES AT LOW SPEEDS 2 Cimer etice zal 8 Sena sree “ : beth a8") O18 lo “179° ot 2a s 9 @ ase a=224 | |e 9-181" 49] 324 a s LP fe ayazerr, 9273 | Fie =201% 0 2 4 60-80 1000 20-40 €o 80 160 Peren char 36-8 5 ro 51.06 a REE Ho dos al Iso ot cy, aeg Feu 718 24] 140° Jos % ao) 0 oz t6| fo or rl ° 1216 20 20 ae HO ay deg Section characteristics; flagged symbol indicates leading-edge roughness. Fraume 06,—Blade-surfaco pressure distributions and blade section characteristics for tho caseade combination, 660°, 1.50, and blede section, NACA 65-(12)10, aioe ool | ti : Kh 4 jars eor 0m" | fia ay : A a ee aor ae| faa, 200 tar ad sq w s ae | 4 [ie «, 240%, 8 te, 290%; 8 305" oe to 8G ao a Ob eS Se _— co 308 eta 32} ee {50 04 Uy a = al io tos,, aden | Se $4 ond 2a 4 130 loz “a zal on ‘LT ' 16 20, 2a 28" ° io ) Section characterstice; Gagged aymbol indicates leading-edge roughness. Fiauns 67.—Blade-eurface pressure distributions and blade section charsoteristics for tho eascade combination, f= 60°, o=1.50, and blade seotion, NACA 05-(15)10. 157758 ‘REPORT 1368—NATIONAL ADVISORY COMMITTED FOR AERONAUTICS Soret sate (0) a «190%; 8-298" WO) 9, = 220%; 0232.2" rT 2al fae - r s 7 [hg es = 2.4] (2) a1=250%; 8) s rT-] ot +4} HH |} il @ @ ar2agr, OSB d 4a —| ae - Eee Li} || (a= 300%; O38. sq sat eat r > Re =| r] s ‘ 1.6} a in anton, aka | | 3 0 845-0 80 WO 0-245 eGo bt oo ‘Percent chord aE 4 70 907 oe 004.10 aol al 0 Jos stk 0) 4 > io ei elea wo Jos a: | ee i er wl | lao Jose, ak 7 © foe, aon Le fe . 7 A) 50-4405 and aol sl 30° Joss 2, 609 36h + ose xt 40 doe zal al 20 Joe at oe at 3 10 for af 3 2 Joe ut 20} 2 10 401 a a a ae lence ae 16 20 24 2 32 oc) @) Section characteristics; fogged symbol indicates leading-edge roughness. Ficuur 68.—Blade-surface pressure distributions and blade section characteristic for the cascade combination, f:—=80°, 4=1.50, and blade wetion, NACA 65-(8)10, (© Seotion charnotoraties, Miovan 09—Blade-surfaco pressure distributions and blade ‘section characteristics for tho eascade combination, ®.=00 4=1.50, and blade section, NACA 65-(23)10.SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIES COMPRESSOR BLADES ATT LOW SPREDS @ar230r, 06349" | [i 92250; 0=380° 2a 3 ches secs sel 8G co EN 32] s vg 4 ee Pecnt So aor 7 er @ 198 ta of HPS wl | FL bao ow ag 20-44) 1% 303-403 0% a > doe y 10 {or ed [a a deg (© Section characterise; ged symbol indicates lading ‘igo routine. Fravne 70—Bladesurface pressure distributions and blad ‘seetion characteristics for the cascade eombinstion, 00", 171.80, and blade seotion, NACA 65-(24)10. 759 2 Cont aaiice A 8 Sora a 6 5 pppetes af o -24" 0.4 Qj 2al t s ‘ a (@ 9240", O15" (6) 4760" 0-3." 4d 32| 2d s 1g a = a) a)=9.0%; 0 =5.0° ee Para chon ar 5 160 407 ar | so to6 le 3} 140 los ab | 10 04, ado by 4 ao 120 403“! apo to foe oF xt © jo oa a 3 en ceg (Section characteristics; flagged symbol indicates lending edge roughness; solid symbol Indieates high Reynolds ‘number. Frovns 71.—Bladesurface pressuro distributions and blade seotion charastristies for the cascade combination, =70°, 1,00, end blade section, NACA 65-010.760 REPORT 1368—NATIONAL ADVISORY COMMITTED FOR AERONAUTICS 2a a Cit arcs 2a a. : | a C a eS Fa N ! A * aye 72; 0-86 id) ays9.2; 0410.8 a _| . ad Co . : ad : is s KI “pes : PRECH | i fh ayst5.27; 812.0" 165; 8-103 0 lor ra @ Section charscteristicn; Gagged symbol indicates leading edge roughness; aolid' symbol indicates high Reynolds number Ficums 72,—Blade-surface pressure distributions and blade ‘ection characteristis for the enseade combination, y= 70", 71.00, and bade section, NACA 85-410, 0 20-40-€0-80 105-920 40-60 80-100 Para chord 70 08 co Jor 2 fos 10 705 cay 47) ond 130° 04 Sr 20 tos 10 toa © or wo to yea (© Seotion characteristics; Aagged symbol indicates lending edge roughness; solid’ symbol indicates high Reynolds number. Figune 78—Blade-surface pressure distributions and blado section characteristia forthe enseade combination, r= 7 100, snd blade section, NACA 85-810,SYSTEMATIC TWO-DIMENSIONAL CASCADM TESTS OF NACA 65-SHRIES COMPRESSOR BLADES AT LOW SPERDS 761 2a} 62r | [er arecr; a-iee al Ms im s (© airl06%: i126 3. 4 a 2 s oe SS @ariasyoeiter | Ro 0 8% & BD 6. 52 25-40-6080 105 a 9 co 308 oa i Bau edi at ol 20 os 16} Ae 40 1.04 “ 489 12-456 20 loser a 4 20 oe + 4 10 dou co j ad 39 ae i216 20° 2 €09 ) Section characteristics; Sugged symbol indicates leading-edge roughness. Frouns 74—Bladesurface preasure distributions and blade ‘ection characteristics for the earcade combination, f= 70°, 1,00, and blade section, NACA 65-(12)10. — ; N , Pst st Pky 7 "a : . ce : . x ad ; tat 08 40 6 80 100-020 a0 80 80 100 Peron chor 2a 60 05 ma lor 3 ie 6° anaes (9 Scotion characteristics; no design point was obtained; ‘Aagged symbol indicates leading-edge roughness; soli symbol indioates high Reynolds number. Froumn 76.—Blade-surtace pressure distributions and blade soction charactartistis for the eatcade combination, :=70°, 21.00, and blade eeotion, NACA 85~(18)10.762 REPORT 1868—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 3 inet oa eal 3 ie 3 Gal ie al Soe ait i s b. . 5 ° | | 2 | ria (0) ay = ee (W) a, 26.0% 8=5.2° — 24) ue s s a g @ ai100% 8-85" 40} 49] ad aa 5 s 2a ug s ps ‘ Jaariamertte | [in w=i67% 8245" of 3 40-6 80 Wo 0B a0 G80 10 i) ay 18.0% O17? Lanne atl 8 707.08 0 & a0 6 80 60 a od ae 108 7 4 60 +1.05 a Aree I a Say | om 6|* 04 4) et 140 16) ¢ is ob aay 8,609 12} 4 9 20 $4.05 0% een bay & al of al so 402 a ao ; 10 ou 3 20 for i id 1 [ee ro Beis eo? ~ ° a a a a deg ay 6 (q) Seotion characteristics; dagged symbol indicates leading: ‘eige roughness; solid symbol indicates high Reynolds ‘number. Provan 76.—Blade-eurface pressure distributions and lade ‘section characteristics for the eateade combination, fm70", 125, and bade section, NACA 65-£10 (@ Section charactoristcs; flagged symbol indicatea lending ‘edge roughness; 2olid symbol indientes high Reynolds umber Fravan 77.—Blade-surfaco prewoure distributions and blade ‘soation charaetariaties for the caseado combination, #}=70", = 125, and blade cection, NACA 65-810.SYSTEMATIC TWO-DISENSIONAL CASCADE THSTS OF NACA 65-SERIBS COMPRESSOR BLADES Av LOW aPEEDS 763, @) ater; 8+ 7 ©) ay 10K"; 8 = Bae 2al 2 Gam eroca 8 Genome astoce ey 56" © arl6h; 8-70" 324 24 i 60-8005 7 Lo4ey9 ond los laf ew ee owes (@ Section eharstaritics; Sagsed symbol SndeatesInding- edge. roughness; anid npmbol indicates high aynolds umber. Figuns 78.—Bledeaurtace preaare dstrbstions and blade seston characteristics for fo casade combination, B= 70%, sH125, and blade section, NAGA 65-(9)0 3 Som ae 2a oan (eh, 10.0% 8 12.0%; 818 324 2al 0-25-40 65 80 105 cent Pecet a ea . oat _ 0. fou at 6 fi so fos oa aes bq, 4 How a st | bo bl 20 for a ° (0 Seotion characteristics; fgged symbol indicates leading fdge roughness; solid ‘symbol indicates high Roynolds number. Fravaz 70—Blade-surfaco pressure distributions and blade seation characteristics for the cascado combination, 8, =70°, 25, and bade section, NACA 65-(15)10.764 REPORT 1308—NANIONAL ADVISORY COMMITTEE FOR AERONAUTICS 3 ies 2. Bee 2 Set aie i hes |_| att a s | | hse (0) 515 00" q 2a \ . Lt I vo te c bet @ 80° : Cl @ aa }-}+-+ a 2a Lt sel-p-tt s p tH 2 PY tt INT ET TY s i F 4 ‘6 © anlar 0 A665 80 eo 9 40-6 Ho Ses ar 4 “0 7.0 @ aj= 40%; 8 1g | [in aye 180% 8 8 03 a0 a 8 Woo ade 80 ToD a 2% Jos Peer cet 29 8 ie 2 Jos ao 4 oan Fao os re 0 Jos oe fa i loa oF] ose ft te 2 Ff _ 4 be es } food Josend ¥' tL % d -2o J on + 4 Ajo -|02 ~ 0 4 o ol oto cao, a) or a Bi ae Jo a 9 (@) Section characteristies; aged symbol indicates Teading- ‘edge roughness; solid aymbol indicates igh Reynolds number. ‘Fravmx 80.—Blade-surfaco pressure distributions and blade section characteristic for the cascade combination, x= 70°, 1.60, aud blade sation, NACA 65-010. % ‘audes @ Section charactersties; Magged symbol indicates leading ego roughness; solid. symbol indicates high ‘Reynolds ‘number. Fioune 1.—Blede-surfaco pressure distributions and blade section characteristics for the cascade combination, 670°, 751.60, and blade seation, NACA 65-410,765 SYSTEMATIC TWO-DIMENSIONAL, CASCADE THSTS OF NAGA 6O-SBRIES COMPRESSOR BLADES AT LOW SPEEDS Cee @) arr; 8-80" 7818." 2 eal iff tod . I te d 2al =| sia | [i oi6ir; Oe197° : Pos Patol. | as on so = +++ eco aa i 3a| at tf} || al s E s HH 1 a a Wariar oer |_fin geo: 0-20-80 80629 40 @ 80 Wo 0B ao B BO WD _—9 BAO BOTS Puce cud od 24 a) 2 60 408 3 3 a co os mt 928i bp 0 fos HEE ig ot ECE tL Te do afm 8, cag jt Nao S408 ot 6, dey 12h. % 4 20 + 20 4o2 Lv \ | | 4 ie 10 for Ag 10 for Fl Ci Zz 7 1% 16 20° y og 12 16 20 2° : udeg 9 () Section charsoterstioa; flagged symbol indicates leading ‘edge. roughness; solid ‘eymbol indioates high Reynolds umber. Frauns 82—Blade-surface pressure distributions and Dado ‘eotion characteristies for the cascade combination, #=70°, ‘7-11.50, aod blade testion, NACA 65-810. @ Section cherastorztos; flagged symbol indicates leading ‘edge roughness; solid “symbol indleatea high Reynolds ‘umber. Fiauns 88—Bladé-eurfae presure distributions and blade ‘svetion characterities forthe cascade embination, @1=70°, 1.60, and blade sootion, NACA 65-(12)10.766 REPORT 1368-—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS aS ame C] : 7 4 RS K janes ter | feecaan tebe al ttt s ‘1 Ps 4 (he ay + 16.0"; 8 20.4" (9) | ay218.0"; Oe ad a |_| od : | “al Ae C] 19) +2007: Oreo? ay 22.0%; B62 0 #6 40 60 80 1090-20 40 60 80 100 Pert cord Me 0 8 © at at aby ot so z0 -foe “ 10 or wa sell, J, Pe eg Section characteristics; Angyed symbol indioates lending edge roughness; solid ‘symbol indicates high Reynolds ‘umber. ‘Provan #4.—Blade-surface pressure distributions and blade ‘sation oharaoteristies for the cascade combination, f.=70°, 1.50, and blade sestion, NACA 05-(15)10. eyo number, 0245x108 moo blose 445x108 smooth lee, txbulanc added [245x108 _Inding-edgeroughrass odd al 20] 40] {oa FS lor x 108 asl los 8,009 % lo4 10s joe e eyo punber, (0245x108 Aa5x108 ede ‘eI Bok BR Ro ae LS eek) 609 (0) Comparison forthe anglo-f-attack range for 6.45%, ¢=1.60. () Comparison for tho anglo-of attack rango for 6,=60%, ¢=1.00. ‘Smooth blade, ‘Fioune 85.—Etfect of Reynolds number on turning angle and ‘drag coellcient of the NACA 65-(12)10 blade section for ‘typleal eascade combinations,SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERINS COMPRESGOR BLADES AT LOW SPEEDS 767 forbuence occed 1s 13 tupulnce ose v9 iz age ee ee oo (© Veviation with R noar design angle of attack. Fioums 85-—Conoluded.768 ‘REPORT 1268-—NATIONAL ADVISORY COMMUTER FOR AERONAUTICS | 2% \ P \ Bice9 i rrr 4 aa (as & = a) iso 4} las SS |r| eo ‘ [oe I J x 2 1216 20-28 Comber. c.g ‘Froure 80.—Variation of estimated operating angle-of ‘with camber forthe inet angles of the tests. : ° 6 29) 39 a0 50 Tring angle eg ‘Froonz 67.—Veriation of ideal and test dynaralo-preasur rato across ‘the cascade.52; a : al 1 = | 48) 2 a so t aa 2 s 2et + — + i 32} | 3 aa a f ged a 3 = [4 [ i 2a - a: 60. a 20) — A 20] ia . ie “P 45|_ 4 | rey — 12 . [+] 14 = tof 5 [> yy ia a iT). fT Ot ae 3a as d om a sot « avas 88 —Convanio dat sowing ha atone eatrng and mann dyn pre saan uniog—-Fiouan $2-Veration af hm odng pee ‘angle and inlet angle. rameter, (cen) me With solidity for Inleb ‘angles of 60° and 70°. SAHEIE MOT LY SACVIE UOSBENIOD SATUN-I9 VOY dO SISKL HAVOSVO TWNOINENUC-OAL OLLVIVEIEAS 692,ee ae ay deg Fravns £0.—Comperison of relationships between turning angle and ang!s of nk or epee porourval ord reas wth thom forte 0 8 @ 6 Be ay eg Frours 01.—Surmmary of relationship between turalng augle @ and angle of attack a: for the blade ‘sootions tested st 8.80", ¢—=1.00, one SOLLAVNOWSY NOX HRLLUWNOO IHOSTAGY TYNOILYN—SORT LLOLME— , , , wl [tI [ E oer 35 + LAs ~ oa | ai zg |. ze 5, | rd “| Yi 16) LA " T 3 ea ae Loy CCCr Leo IR ERE | | 1 T 7. 12 20 B eas al Piovny 92.—Summary of relatlonahtp betwoen turning angie @ and angle of attack a for ‘tho blade soctions tasted at fi=30°, em=l.26. ‘uns 03.—Summary of wlstionahip botween turning at ‘tack a for the blade oestons tested at 6 SCUEIS AOT LY SHAVIE WOSVRUAROO SUIEE-GO VOVN £0 RISK LUVOBVO TYNOTENUAIG-O Atul OLLYINGIEES Taki ; + 2 | { a 2a eb at a aot | 4 A al Lat + q + 28 [ a | 2 65-410 4 | 2 (|? Bite 1 1 : tT T |_| ipl pi iii iii | fifi | 1 Fioune 06.—Summary of relationship between turing ‘attack oy for tho blade eeetlone taated at Bim 4B", 9 and angle of -_‘Frawas 05.—Suramary of relationship botwean turning angle @ and angio of sttack for 50. ‘the blade sootions tarted at Bj=48", 20.75. SORAVNOWRY NOK SULIEFOLOO AUOSTAGY TWNONLYN—808T THIOAR49 44] Tr] “ ] T T * I | | Deen, $7 tl | < +4 x L ae a8 wv T I | | zat 2a} F L. t + +t = a T 7 i i i _— \g F | i 1 fo Ad Lt fale : L |: Si. 3 1 TL? ssetile el S 2 H 4 + : 4 AY T T _ 9 -s| 5 T - Be - i 1 2 oe Be 3 onde a doo ‘Fravrn 08.—Summary of relationship botwocn turning angle # and angle of atteak ay or the bade sotions tested at p= 48", o=1.00-4 Poovns 97.—Suramary of relationship between turning angle @ and angle ‘of attonk ay for tho blado septions tested at f= 6", ¢=1.2, SNAG MOT LY SHAVTH HOBERUENOO SHNNE-I9 YOVN £O SIBEL MAVOSVD TVNOISNEATE-OAML OLVREIAAS ele[ | I T “ [TT - j { _{— Cy sat “CPT | | T T } | wilt A eel) rT | i TT | 7 Coo CAA “ : T 7 7 ml “co COP eT ‘ 7 lala) alalalelal b OO r =COCOCOE eee al LT z t 3 { | | 4 16 as 609 Fraoas 08.—Sunmary of relationship bobween turalng angle ¢ and angle of attack for ‘the blade sootions tested at pi=d6°, ©= 1.60. 7 i % en deg Frovan 99.—Summary of relationship between, turning angle and angle of attack forthe blade eoetions tasted at 6,0", «0.50. PLL SOMLAYNOURY WO GALLWNOD AUOSLAGY IVNOMLYN—808L SHOEsooo L000 J Pay T poems] | { i T - 2 nal | TF rT 20} 20) ~ | {__ | Lo | _| | i is Y | 1 + a —- 2 a i | lel | i ps cal 4 : xO | a r | = oa eee | a | ee i ; 6- | I 4 2 jean > 6. | Wa Avi | Ly, | | | 7 } “* " se 16 2 a 4 S 16 ay Frauma 100—Bummery of relationship bebwoen turning angle @and engloo! atiaok ay for the blade sectlons tested at B= 00", =0.76. eg Frovan 101.—Summary of relationship botwoan turning angle # and anglo of attask ai for ‘the Blade sections tested at 2-00", e=.00. SCHGSE MOT LY SHAVIE HOSEENANOD SATME-S9 VOVN AO SISKL AVOSYD TVNOISNEREC-OML OTLYREELAAS aLa4 44 1 T +0} 7 Dain x | 38| 3 . 32 Det 2 2e| a 2 A | Jt4 E gL 3 a i aa i | 2 - ' © 6500 q 4 © ea & E810 2 S20 - 2 erti3i0 ! : ges 4 2 eat aA 3 Slash o od -al | l 7 L 1 LLL a. 4 rn a a ay des : ay deg Frau 102.—Summery of rdationship between taming angle $ and ‘angle of attack ay forthe blade eections tested at =80", om1.23. Fiovas 108. —Summary of relationship bebroen turning angle @ and angle of attack a for tho blade sections tested st 600%, = 1.50. nk SOMAVNOWMY UO ESLITANOO KNOBLAGY TYNOLLYN—892T THOMEx 20) . ro if oer | oh $4 i uy t aS i £ i a L = = . | ; - c 2a 3 $48 cae: 3 Sa $ Bee 2 BiB. |] 3s (1210 & eis) o| o| -4 ag Fraoaa 104—Summary of ralatlonahip between turning angle @ and angle of attack ay for ag ‘ho blade seotious tasted at 10%, 1.00. Provme 105.—Suinmary of rolatlonship between turng angle @ and angle of 5 7 7 oy dog attack ay for the blade eootions bated at &;=70°, 01.25, SOMES ALOT LY SHOVIE HOREAMENOD SUINES-I9 YOVN ZO SIMUL REVOSVD IVNOIGNGHIG-OM OMYRELARS bbb106.—Suromary of relationship between tring angle @ and angle of attack: a for the ‘lade sootions tested at ge70°, #=1.50. © §-00 8 eeeaio | 2 eI 2p | 2 eetig}i0 8 eEUshO | T 8 72 16 24 ° 2 Te 18 ay 08 Saidity, « roves 107.—Varfation of slope of curves of taming angle aginst angle of attack at design conditions with solidly ‘and inlet angle, The slopes are averages for the moderato camber Fangs. BLL. SOLAYNOWAY HOR SRLIZONOD ANOSTAGY TYNOTLYN—80ET GAOT,BE 7 po rT ] i - | 2a _ t — 65-1240 a - tpt SSAETIO io ‘isn _ : (150 t6}-— t [= BIBI [Sot a SA A + BIE = | = 65-410 a aed 1 i—} | || d es-010 ‘ | [ f ——— [+7 t—~_| | ° « 5 5 io 12 i# i is Salty, & ‘Frovnn 108,—Varltlon of design anglo of attask with solldlty for the sections teste, SORHIS MOT LY SuAVTE HOSSELENOD HINEE-I9 VOVN £0 SISK RGVOSTD TNOINERTC-OMLL OLLYHRLESS LLDesign turing ong y Soidhy,& () Tolet angles of 60° and 70°, Variation of design turning angle with solidity and inlet anglo for tho ections tested. SOLLOVNONSY NOM HELIAONOO HOSTACY TVNOLLYN—SogT waa 08.‘SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA O5-SERINS COMPRESSOR BLADES AT LOW smDS 781 4 p-—| 160. B12 aq 2 eh ks-t810 2 mp 3 — ls-tano ;-—~ ct] | [+] 8-410 a6 39 5 70 30 Entering o ona, ‘Frovnn 110,—Variation of design turning angle with inet anglo and solidity for typieal sections,REPOR? 1365—NATIONAL ADVISORY COMMITTEE FOR ABRONAUTICS Bese 30 a Conte G5 @) e=o7s © «1.00. Fravan 111.—Variation of design turning angle and design angle of attack with eamber and inlet angle.SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SHRIES COMPRESSOR BLADES AT LOW SPEEDS 783 3 a3 e : a 7 : i a Ae = 7 a : “ ie = 1» lz wh oP 2 — 4 Paco oF 24 + £ | 7o| B x it S I 1 & T F - | | wo T <7 5 =e 7 7 aie os os Fravas 111.—Coneluded. 20070002REPORT 1969—NATIONAL ADVISORY COMAIITER FOR AERONAUTICS Frovaz 112.—Design-camber soleation chart for NACA 65-(¢,.)10 sectionSYSTEMATIC TWO-DIMENSIONAL. CASCADE TESTS OF NACA 65-SERIES COMPRESSOR BLADES AT LOW SPEEDS 785 ‘igump 112—Coneluded,786 [REPORT 1368—NATIONAL ADVISORY COMMINTEB YOR AERONAUTICS 2 2 cr 2 f ‘ a r \ i f . 1 0 a_| by et o 4 S y cA eI A es ° VV | A Fiaune 113—Design angle-of-attack chart for NACA 65-(ex,)10 sectionsSYSYBMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA G5-SERIES COMPRESSOR BLADES A'? LOW SPEEDS i anh 1 609 Frovns 113.—Coneluded, 787788 REPORT 1563—NATIONAL ADVISORY COMMITTBE FOR AERONAUTICS ‘Fiouns 114.—Otf-design turning-angle carpets for NACA 65-(c,,)10 compressor blades,SYSYDMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA 65-SERIES COMPRESSOR BLADES AT LOW SPEEDS 789 cia Fioons 114,—Continued,790 REPORT 1808—NATIONAL. ADVISORY COMACTTER FOR AERONAUTICS ae ~ — HH LH aH | 99 aa | T 95 ct 91 3e/+ t 1 « fo corr $5582, r 7 16 or ia 79} Py al L cla] 2a £ ¥, 8, 60 56 a+ 212 te 1 | { { rooms 114.—Continued.SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS OF NACA G5-SERIES COMPRESSOR BLADES AT LOW SPEEDS 791 99 44 . 103 48 Ey el I Ae a1 36 + + a 2 vdeo 0, 600 | 79 ab ms nm 16 er i 6 a ol a 0 4o|—}—} T 7% 3 a, deg 192 ee a aoa ae Froune 114.—Continued.792, na No 02 we m 70 46; REPORT 1368—NATIONAL ADVISORY COMMOTTEE FOR AERONAUTICS Nk ava 8 Be | se 1 3 }—| + es cc) ‘Fiavan 114.—Continued.SYSTEMATIC TWO-DIMENSIONAL CASCADE THSTS OP NACA O5-SERIES COMPRUSSOR BLADES AT LOW SPEEDS 793. a 4 naa 108 = woe “2 _ 98 {| 8 2568 24 es eu eo. 3 | m 4 @ 7m ot Fravrs 114—Coneluded.REPORT 1363—NATIONAL ADVISORY COMMIPTEE FOR AERONAUTICS 3 = 4 2 35 a tae By, ses pce Bude Bude ‘Fraunn 115.—Composite of of-desien carpets for NACA 65-serlessootions for values of ex. of 0.4, 1.2, and 1.8,SYSTEMATIC TWO-DIMGINSIONAL CASCADE TESTS OF NACA 05-SERIBS COMPRESSOR BLADES AT LOW SPEEDS 795 Figuns 115.—Coneluded.796 REPORT 1368—NATIONAL ADVISORY COMMITEE FOR AERONAUTICS ARERR SEN RRR TS apes ah Bel a4 0.50 () Constant stream-tubo area. {() Constant axial velocity. Frovme 116.—Prearure-isecooisiont carpets.SYSTEMATIC TWO-DIMDINSIONAL CASCADE THSTS OF NACA 65-SERIES COMPRESSOR BLADES AT LOW SPEEDS 797 (@) Gonotuded. (©) Concludes. Fravs 16—Conoluded.