What was the purpose of the tests conducted in the NASA Ames Eleven-Foot Transonic Wind Tunnel?

The tests aimed to evaluate advanced rotorcraft airfoils to confirm predicted performance and validate advanced airfoil design methods.

What airfoils were tested in the wind tunnel?

The tested airfoils included two modern baseline airfoils, SC1095 and SC1094 R8, and three advanced transonic airfoils, SSC-A09, SSC-A07, and SSC-B08.

What were the maximum lift coefficients observed during the tests?

The maximum lift coefficients observed were 1.37 for SC1095, 1.72 for SC1094 R8, and 1.40, 1.22, and 1.15 for the SSC-A09, SSC-A07, and SSC-B08 airfoils, respectively.

How does the performance of the tested airfoils compare to previous tests?

The maximum lift coefficients for SC1095 and SC1094 R8 were about 9% above prior test values, indicating improved performance.

What is the significance of the Eleven-Foot Transonic Wind Tunnel for airfoil testing?

The Eleven-Foot Transonic Wind Tunnel is significant as it can operate at a wide range of Mach numbers and provides a suitable environment for two-dimensional airfoil testing.

An Experimental Evaluation of Advanced Rotorcraft Airfoils in the NASA Ames Eleven-foot Transonic Wind Tunnel

Document

166587 NASA CONTRACTOR REPORT An Experimental Evaluation of Advanced Rotorcraft Airfoils in the NASA Ames Eleven-Foot Transonic Wind Tunnel Robert J. Flemming _,t R_'3- 1 !(O,n (tlASA-CF-166587) AN EXPEPIMENT_L EVALU_ rIOl!

OF ADVANCED ROTOECRKFT AIFFOIL_ i'J THe: N_5_ AMES ELEVEII-FOOT TRA_ISONNC I:IND TU_'NEL Contractor Pepoc%, P.ar. 1982 - _pr_ 1983 GBIC2 :Sikorsky Aircraft) 162 p Avail: NTIS HC CONTRACT MOA 14800-039 September 1984

NASA

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J a, NASA CONTRACTOR REPORT 166587 An Experimental Evaluation of Advanced Rotorcraft Airfoils in The NASA Ames Eleven-Foot Wind Tunnel Robert J. Flemming United Technologies Sikorsky Aircraft Prepared for Ames Research Center under Contract M0A 14800-039

N6SA

National Aeronautics and Space Administration Ames Research Center Moffett Field, California 94035 _.1 1!

FOREWORD The test and data comparisons contained in this report are the result of a cooperative rotorcraft airfoil program between the Sikorsky Aircraft Division of United Technologies Corporation and the Ames Research Center of the National Aeronautics and Space Administration. While the tested airfoils are the product of Sikorsky design efforts, the test data and theoret- ical comparisons are published herein to advance the state of rotorcraft airfoil performance pred/ction. Several comparisons are contained in this report, but the reader is invited to use the data to provide additional insight into the areas where the available theoretical methods give valid results and where further theory development is required.

Many people provided the technical support to conduct this pro- gram. The principal personnel include: NASA Ames Raymond Hicks Project Coordination LeRoy Guist NASA Ames Test Engineer Donald Jepson Sikorsky Aircraft Model Design Anthony Saccullo Sikorsky Aircraft Test Engineer David Lednicer Sikorsky Aircraft Aerodynamicist PI_CE_NG PAGE. BIJ_K NOT Fit..M:-"D iii TABLE OF •CONTENTS Foreword iii List of Tables V List of Figures vi Summary i Introduction Symbols Test Facility Models Instrumentation Test Procedure Data Reduction Methods Test Results Theory Correlation i0 Conclusions Appendix A - Tabulated Test Data References iv LIST OF TABLES I Airfoil Model Geometric Characteristics i3 II Coordinates for the SC1095 and SC1094 R8 Airfoils 13 III Model Configuration Summary 15 IV Run Log i_ V Estimated Data Accuracy 25 V LIST OF FIGURES I.

TSA installed in the Ames Eleven-Foot Transonic Wind Tunnel .

TSA schematic 3.

Airfoil section profiles 4.

Airfoil metric sections 5.

Test Reynolds numbers 6.

Metric section calibration fixture 7.

Representative manometer board wake rake profiles 31 8.

Data repeatability - SCi095 airfoil, Mach number = 0.40 a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient , Balance measurement correlation with pressure 35 measurements a. Lift b. Drag c. Pitching moment Drag for a drag coefficient of 0.008 Aerodynamic characteristics of the SCI095 airfoil a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient 12.

Aerodynamic characteristics of the SSC-A09 airfoil a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient 13.

Aerodynamic characteristics of the SSC-A07 airfoil a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient vi !_! I i LIST OF FIGURES (Cont'd) 14.

Aerodynamlc characteristics of the SSC-B08 airfoil a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient 15.

Aerodynamlc characteristics of the SC1094 R8 airfoil a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient 16.

Aerodynamlc characteristics at a Mach number of 0.30 ao Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient 17.

Aerodynamlc characteristics at a Mach number of 0.40 a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient

18. 6O

Aerodynamlc characteristics at a Mach number of 0.50 a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient 19.

Aerodynamlc characteristics at a Mach number of 0.60 63 a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient 20.

Aerodynamlc characteristics at a Mach number of 0.70 66 a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient 21.

Aerodynamlc characteristics at a Mach number of 0.80 6g a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient vii LIST OF FIGURES (Cont'd) 22.

Aerodynamic characteristics at a Mach number of 0.85 a. Lift coefficient versus angle of attack b. Drag coefficient versus lift coefficient c. Pitching moment coefficient versus lift coefficient 23.

Variation in maximum lift coefficient versus Mach number 24.

Variation in drag coefficient at zero lift versus Mach number 25.

Maximum L/D versus Mach number 26.

Lift curve slope correlation 27.

Pressure coefficient distribution for the SCi095 airfoil a. M = 0.40 b. M = 0.60 c. M = 0.80 28.

Pressure coefficient distribution for the SSC-A09 airfoil a. M = 0.40 b. M = 0.60 C. M = 0.80 29.

Pressure coefficient distribution for the SSC-A07 airfoil a. M = 0.40 b. M = 0.60 c. M = 0.80 30.

Pressure coefficient distribution for the SSC-B08 airfoil a. M = 0.40 b. M = 0.60 c. M = 0.80 31.

Pressure coefficient distribution for the SCI094 R8 91 airfoil a. M = 0.40 b. M = 0.60 c. M = 0.80 viii _!If LIST OF FIGL_ES (Ccnt'd) 32. Pressure coefficient distribution for low lift at high Mach numbers a. M = .825 b. M - .85 c. M = .88 d. M = .90 e. M = .98 f. M= 1.07 i00 33.

Pressure coefficient correlation, M = 0.30, C 1 = 0 a. SCI095 b. SSC-A09 c. SSC-A07 d. SSC-B08 e. SCI094 R8 34.

Pressure coefficient correlation, M = 0.30, C 1 = 1.2 a. SC1095 b. SSC-A09 c. SSC-A07 d. SSC-B08 e. SCI094 R8 f. SCI094 R8, C 1 = 1.5 35. iii Pressure coefficient correlation, M = 0.4, C 1 = .7 a. SCI095 b. SSC-A09 c. SSC-A07 d. SSC-B08 e. SCI094 R8 36. 116 Pressure coefficient correlation, M = 0.6, C 1 = .4 a. SCI095 b. SSC-A09 c. SSC-A07 d. SSC-B08 e. SCI094 R8 37.

Pressure coefficient correlation, M = 0.825, C 1 = 0 a. SC1095 b. SSC-A09 c. SSC-A07 d. SSC-B08 e. SC1094 R8 ix _'I 11 An Experimental Evaluation of Advanced Rotorcraft Airfoils in the NASA Ames Eleven-Foot Transonic Wind Tunnel R. J. Flemming Sikorsky Aircraft SUDI4ARY Five full scale rotorcraft airfoils were tested in March and April 1982 in the NASA Ames Eleven-Foot Transonic Wind Tunnel for full scale Reynolds numbers at Mach numbers from 0.3 to 1.07. The models, which spanned the tunnel from floor to ceiling, included two modern baseline airfoils, the SC1095 and SCI094 RS, which have been previously tested in other facil- ities. Three advanced transonic airfoils, designated the SSC-A09, SSC-A07, and SSC-B08, were tested to confirm predicted performance and provide confirmation of advanced airfoil design methods. This test has shown that the eleven-foot tunnel is suited to two-dimensional airfoil testing.

The maximum lift coefficients at a Mach number of 0.3 for the SC1095 and SC1094 R8 were 1.37 and 1.72, respectively, about 9% above prior test values. The transonic airfoils had maximum lift coefficients of 1.40, 1.22, and 1.15 for the SSC-A09, -B08 and -A07, respectively. Drag divergence Mach numbers at zero lift for these airfoils were .808, .780, .833, .848 and .860.

Prior to stall and drag divergence the pitching moments were generally between 0.010 and -0.015. SC1095 and $CI094 R8 lift curve slopes were 8 to 17_ below that of the solid-wall United Technologies Research Center tunnel, used to test the baseline airfoils in 1975.

The airfoil analysis codes agreed well with this data, with the Grumman GRUMFOIL code giving the best overall performance correlation. The NYU Transonic Airfoil code predicted airfoil pressures and drag divergence well, but errs in the calculation of pitching moment. The Texas A&M TRANDES/TRANSEP codes show good correlation over the full range of test conditions. The AMI CLMAX code predicts the relative maximum lift coefficient of the thicker airfoils well, but fails to predict the maximum lift coefficient of the S$C-A07. The maximum lift coefficients measured in the test exceed the CLMAX code prediction and available test data from the United Technologies tunnel by about 10%.

INTRODUCT ION Rotor systems must be improved to satisfy mission requirements which demand advancements in efficiency for higher cruise speeds and lower fuel consumption and for reductions in acous- tic levels. Advances in methodology have provided more rig- orous means to design improved airfoils, but these codes have not had a good correlation base for rotorcraft airfoils - airfoils that have compromises between high lift at low velo- cities and low drag at transonic velocities, all while main- raining low pitching moments.

Sikorsky Aircraft initiated a project in 1979 to replace the SC1095 airfoil family with a family of airfoils that maintain its maximum lift capability and pitching moment levels while increasing drag divergence Mach number by .03 or more. This airfoil family was designated the SSC-AXX family. An addi- tional design incorporated a different design philosophy to provide a pitching moment near zero. This airfoil family was designated the SSC-BXX family. The design study used many airfoil codes, including TRANDES, NYU Transonic code (program H), AMI's CLMAX code, FIX) 6, and GRUMFOIL (MCMJ-9) (refs. i-5).

While these codes correlate well with modern airfoils such as the SCi095, additional data was required to validate the new transonic airfoil designs and the theories that were used to design them. A cooperative two-dimensional test program between NASA's Ames Research Center and Sikorsky Aircraft was initiated in 1980 to satisfy these validation requirements.

This report describes the test procedure, data analysis methods, processed data, and code correlation for this test program, conducted in the Ames Eleven-Foot Transonic Wind Tunnel.

I II [!

SYMBOLS A Axial Force, kg (lb) C Airfoil Chord, m (ft) Axial Force Coefficient, A/Sq

cA

Drag Coefficient, D/Sg C d Lift Coefficient, L/Sq C 1 Pitching moment coefficient reference to C m quarter chord, PM/Scq Normal Force Coefficient, N/Sg CN Surface Pressure Coefficient, PI-P- Cp

%

D Drag, newtons ( ib ) h Tunnel height, m (ft) L Lift, newtons (Ib ) M Mach number Mach number for drag divergence, dCd/dM = 0.1 N Normal Force, newtons ( lb ) P Pressure, newtons/m 2 (psf) PM Pitching Moment, newton-m (ft-lb) g Dynamic pressure, _oV 2, newtons/m 2 (psf) Reynolds Number

N

S Metric Section Area, m 2 (ft 2) t Airfoil Thickness, cm (in) V Velocity, mps (fps) Angle of Attack, deg p Air density, newtons/m 3, (slugs/ft 3) Subscripts BAL Balance 1 Local max Maximum P Pressure w Wake Free Stream TEST FACILITY The Eleven-Foot Transonic Wind Tunnel at NASA Ames is part of the Unitary Plan Wind Tunnel complex. It is a closed return, variable density tunnel with airflow produced by a r2Lree-stage axial-flow compressor. The tunnel can be operated at Mach numbers from 0.4 to 1.4 at stagnation pressures from 0.5 to 2.25 atmospheres and at lower Mach numbers at pressures above 1.4 atmospheres. For the advanced rotorcraft airfoil test the maximum Mach number was 1.07 and the stagnation pressure was held at 1.0 and 1.4 atmospheres. Stagnation temperature averaged 294°K (530OR).

The four walls of the test section are slotted with a normal porosity of 5.1%. To provide smooth flow near the ends of the airfoil model the slots adjacent to the model were taped, reducing porosity to 4.7_.

MODELS The Sikorsky Tunnel Spanning Apparatus (TSA) was instal!ed in the eleven-foot tunnel in a vertical orientation (see fig. 1).

Dimensional data for the TSA is provided in figure 2. The base of the TSA's stainless steel spar was adapted to the tunnel yaw table and a turntable was fabricated to support the upper end of the spar. The turntables were controlled by one primary input with trim adjustments made with the upper turntable controller. Seven fiberglass-graphite airfoil panel segments for each airfoil model were attached to the spar. Surface pressures were measured using 24 upper surface and 11 lower surface .107 cm (.042 inch) orifices located i5.24 cm (6 inches) above the model centerline. The center 20.32 cm (8 inches) of the model contains a six-component Task balance and a single-component rear load cell. The metric section is sealed to the non-metric panels with .024 cm (.010 in) thick elastomeric material. Two struts with triangular cross-sec- tions provided part-span support. The test of Reference 6 showed that the struts do not affect airfoil performance.

Five airfoil profiles (fig. 3) were fabricated for this test, including the SC1095 and SC1094 R8 for which test data in other facilities was already available. The chords of these two models are about 41 cm (16 inches). The three advanced airfoil models fabricated for this test have chords of 43.9 to 54.2 cm (17.3 to 21.3 inches). The chord increase was required to accomodate the spar for these airfoils, which are thinner than the SC1095. The airfoil metric sections are shown in figure 4.

Tests near atmospheric pressure provide full scale data for aircraft in the size range of the Sikorsky S-76 and UH-60A, Bell UR-1H, and Hughes AH-64A.

While the tunnel can be operated over a wide range of stagna-

tion pressures, data were acquired at pressures of 76 ca (30 inches) and 107 cm (42 inches) of mercury. The latter pressure was required at M = 0.3 because of minimum motor RPM con st.Taints. The SSC-A09 airfoil was operated at Mach numbers up to .84 at both pressures to define Reynolds number trends. The test Reynolds numbers are summarized in figure 5.

Table I shows the basic geometric properties of the airfoil models. The coordinates for the SCI095 and SCI094 R8 airfoil sections are given on the first page of Table I I. The coordi- nates for the SSC-A09 and SSC-A07 sections for which a patent is pending and the SSC-B08 section are included on the second page of Table If. The airfoil section profiles (fig. 3) were produced from aluminum molds using fiberglass with stiffening provided by graphite strips. This fabrication process general- ly produced airfoils to a tolerance within .03 cm (.012 inches). The panel segments of the SSC-A09 airfoil were reworked prior to Run lg6 to reduce bolt head loads. This re- sulted in larger tolerance errors Comparison of data taken prior to the modification with that after the modification indicates that the data of Runs 196-221 has a reduction in C_m=T of 0.11, an increase in drag of 0.0014 and an increase in pIT_l_ing moment of .001. This is discussed further later in this report (see pageS). Surface finish was smooth, compar- able to production blade finishes Boundary layer transition devices were not used because full scale Reynolds numbers were used during testing.

At the end of the test, several out-of-contour modifications were made to the SSC-A09 airfoil using tape and wax. The description of these changes is given in the Test Results section of this report.

INSTRUMENTATION The airfoil section forces and moments were derived from the balance readings and by pressure integrations. The center 20.3 cm (8 inches) of the TSA span is mounted to a 2.54 cm (one- inch) diameter six component Task balance and a single com- ponent load cell (see fig. 2). calibration of this system was made with elastomeric seals in place, using special calibration fixtures (fig.6). The balance system was check loaded for each configuration during the test.

Pressures from the model orifices and the sting-mounted wake rake were measured by an automatic scanning system with preci- sion transducers. Half of the wake rake tubes were teed to a mercu_l manometer board to aid in visualization and rake placement (fig. 7).

Model incidence was measured with potentiometers on both the

lower and upper turntables. The TSA spar and struts were strain gauged to permit monitoring of the component loads. All parameters were displayed on digital voltmeters to permit continuous monitoring of the data. Data were recorded on the tunnel data system and transmitted to the Ames computer for on-line data reduction and stored for final post-test process- ing. Final data tapes were transmitted to Sikorsky Aircraft for preparation of final data listings and to facilitate the plotting of data.

TEST PROCEDURZ The test was conducted according to a test plan which pre- scribed angle of attack variation from -5 degrees through stall for Mach numbers between 0.3 and 1.07, except when limited by strut compression loads. Drag divergence Mach number was defined by a Mach number sweep at zero lift. The wake rake was generally covered at Mach numbers of 0.9 and above to prevent vibratory damage to the rake tubes. Ultra- violet oil flow photographs were taken for selected conditions.

Each data point was approached from a lower angle of attack with 30 seconds allowed for the tunnel and manometer board to stabilize prior to data acquisition. Data repeatibility with angle of attack set in both the increasing and decreasing directions was evaluated during the initial test runs. Repeat- ibility is excellent and there are no signs of hysterisis in any parameter (fig. 8). Test repeatibility was checked during each run by repeating the Mach number of 0.4 case at angles of attack of 0 and 6 degrees.

Run The configurations tested are summarized in Table II!.

conditions are presented in Table IV.

DATA REDUCTION METHODS The equations used to transform raw test data to aerodynamic coefficient follow accepted procedures. A description of the equations used in the data reduction process are given below to assist the reader in understanding the derivations of the coefficients.

The aerodynamic parameters contained in this report are cor- rected for the effect of the tunnel walls and spar torsion.

The magn/tude of the wall corrections that must be applied to the data are small. Since airfoil thickness ratios are 9.5_ or less and height to chord ratios greater than 6, the wall correction factors increase the free stream Mach number by l_, the lift and drag coefficients decrease by i_, with small changes to pitching moment and angle of attack. The relation- ships used are given in Reference 7. An additional correction is made to the angle of attack to account for the change in angle at the metric section due to torsional moments. This correction increases the magnitude of the angle of attack about 2%. The lift curve slope in a slotted tunnel is less than that of a solid wall tunnel by 8 to 17_. The angles in this report are not corrected for the slot effect, but corrections are pre- sente--_-in the Test Results section of this report.

The coefficients of lift and drag are presented in the wind axis system. The wake rake drag is measured in the wind axis system, but the balance chord force and balance and surface pressure normal forces must be transformed as follows: C_ = ON(COS a + tan a sin a) - CDtanw a

= cos o - sin

BAL = CNBAL sin a + CABAL cos a The pitching moments for all of the airfoils, except the SCI094 R8, are referenced to the quarter chord. The SCI094 R8 pitch- ing moment is referenced to the quarter chord of the SCi095.

The quarter chord moment for the SCI094 R8 is CM = CM - .0025 C L - .015 C D Use of this trau/sformation increases the nose down moment at high lift conditions by .005.

The wake rake data were analyzed following the procedures of Reference 8. Corrections for wall interference and the velo- city gradient across the probes were applied.

TEST RESULTS The airfoil surface pressure data, internal balance data, and wake rake pressure data were used to produce coefficients of lift, drag and quarter chord pitching moment, presented in tabular form in Appendix A. At low tunnel speeds the coeffi- cients based on pressure 'data are .inherently more accurate.

Model flexibility results in errors an the transfer of loads to the balance, especially in the axial direction. As the tunnel speed is increased, and loads increase the agreement between pressure and balance measurements improve. At high Mach numbers the balance provides more accurate results, since the balance is not affected by force and moment pressure integra- tion uncertainties due to rotational flow and shock position location between pressure ports. A comparison of force and moment coefficients derlved from pressure and balance measure mentsis shown in Figure 9. The lift coefficient agreement is very good, even for cases with shock waves and for post-stall conditions (see fig. 9a). The estimated data accuracy for these measurements is given in Table V.

The wake rake provided much better drag coefficient repeat- ability than the balance measurements. The drag uncertainty for the balance was about 1.5 kilograms due to the flexibility in bond joints between the composite model skins and the balance clamps. (Future metric sections will be machined from solid metal to avoid this flexibility.) This 1.5 kilogram uncertainty exceeds the nominal minimum drag coefficient for Mach numbers below 0.64 (see fig. 10). Figure 9b shows the data scatter that exists in balance drag measurements. While points showing good agreement exist within the overall data scatter, balance drag values for points where the measured wake rake drag is less than 1.5 kilograms are generally not pre- sented in Appendix A. The agreement between balance and pressure-derived pitching moment coefficients are good, improv- ing with increasing Mach number. The plotted data presented in figures 11 through 25 are based on pressure measurements.

Figure 11 shows the force and moment coefficient data for the SC1095 airfoil for a range of Mach numbers. The maximum lift coefficient for the SC1095 at low Mach numbers as measured in the Ames ll-foot wind tunnels exceeds the maximum lift coeffi- cient measured with the TSA in the UTRC 8-foot wind tunnel by 104. Measured drag coefficients agree well. Force and moment coefficient data for the SSC-A09, SSC-A07, SSC-B08, and SC1094 R8 airfoils are presented in figures 12 through 15.

The SSC-A09 airfoil attachment points had to be reworked to reduce bolt head stresses. This resulted in a slight upward rotation of the leading edge piece and a corresponding dis-

continuity between the leading edge and trailing edge parts of

the model for Runs 196 to 285. Post test evaluation of the data showed that this tolerance error caused a degradation in airfoil performance. The drag coefficient increased by 0.0014 and the pitching moment increased by 0.001. The maximum lift coefficient at a Nach number of 0.3 was lower by 0.11 after the rework and the point of zero lift occurs at a 0.3 degree higher angle of attack. Of this block of data only Run 196 is used in the graphical presentations in this report. This run is shown in figure 12 and exhibits a premature reduction in lift coeffi- cient at angles of attack about 13 degrees. The dashed line in figure 12a shows the minimum performance expected for the airfoil at a Mach number of 0.4.

Figures 16 through 22 show the effect of airfoil configuration at constant Mach numbers. Figure 16a shows the low Mach number high lift characteristics of each airfoil. The high lift benefits of the leading edge camber of the SC1094 R8 are evident in this figure. The three transonic airfoils performed satisfactorily at this condition. The SSC-A09 airfoil exceeded the SC1095 airfoil maximum lift coefficient by 2_, and each transonic airfoil tested showed "gentler" stall character- istics. Low lift, low Mach number drag levels ranged from .0067 to .0088. The transonic airfoils had lower drag levels than the baseline airfoils.

The transonic airfoils produced significant performance im- provements at higher Mach numbers. The maximum lift of the SSC-A09 exceeded that of the other airfoils tested at Mach numbers between 0.50 and 0.74. Above a Mach number of 0.74 the SSC-A07 had superior maximum lift capability (see fig. 23).

Figure 24 shows the zero lift drag for the tested airfoils.

The type of leading edge camber used for the Scl0g4 R8 results in an early drag rise and a drag divergence Mach number that is significantly lower than the other airfoils. The transonic airfoils maintain low drag characteristics to Mach numbers above 0.833. The drag divergence Mach number occurs at lower drag levels for the improved airfoils, providing more drag reduction than indicated by changes in drag divergence Mach number. The lift-drag ratios for the airfoils designed using modern design methods are superior to earlier rotorcraft airfoils. The airfoils in the SSC-AXX family have better maximum L/D values than the other tested airfoils (fig. 25).

Slotted wind tunnels give lower lift curve slopes than given in solid wall tunnels or by theory (see ref. g). Figure 26 compares, for the SC1095 and SC1094 R8 airfoils, the lift curve slope derived from theory and the Ames and UTRC tunnels. The differences between tunnels ranges from 8_ at low Mach numbers to 17_ at high Mach numbers.

A limited number of runs at higher Reynolds numbers were made during the latter part of the test. These runs, which were at a Reynolds number 40_ above the baseline, showed little change in maximum lift, a very. small increase in drag coefficient (+.0008), and a small Increase in pitching moment (+.004).

Five types of out-of-contour bumps and protruberances were added to the SSC-A09 airfoil at the end of the test and run over limited angle of attack and Nach number ranges. Each configuration showed a degradation in maximum lift coefficient and an increase in drag coefficient. Pitching moment coeffi- cient changes were generally within ±.005 of the baseline value.

The first change (Configuration 6) was a simulated out-of- contour de-icing boot or abrasion strip. A soft duct tape was .......

applied to the leading edge of the airfoil bac_o an x/c of 10_ for both the upper and lower surfaces. The tape thickness .....

was 0.35_ of chord and ended in a step discontinuity. This .....

resulted in a 15_ reduction in maximum lift and an 80_ increase ........

in drag. This configuration was modified by adding a fairing behind the tape (Configuration 7). The fairings re_ce_ the penalties for configuration 6 by 50_. The effect of miniature _ ......

pressure transducers mounted on the blade surface was investi- gated (configuration 8). Three rows o£ fifteen units, each having a diameter of 0.40 cm and a height of 0.08 cm with a simulated base and wiring, were placed on the model on the pressure orifice line, on the centerline of the metric section and 15 cm below the metric section centerline. The simulated transducers reduced the maximum lift by 4_ and increased the drag by 18_.

Configurations 9 and i0 were smooth surface bumps. The first had a height of 0.3_ of chord centered at the 50_ chord station on the upper surface. The chordwise extent was 29_. This bump caused a 2_ reduction in maximum lift and a 15_ increase in drag. Adding a second bump at 10_ chord (Configuration i0) with a height of 0.2_ of chord and a chordwise extent of 14_ resulted in a further loss in maximum lift of 1_ and a further drag increase of 7_.

THEORY CORRELATION Surface pressure data for the tested airfoils are presented in figures 27-32. These data have been used to compare several analysis methods (figs. 33-37). Figure 33 presents the surface pressure correlation for the five tested airfoils at low lifts and low Mach numbers. The computer codes produced similar results, and match the test data well. Pressure differences near the trailing edge are evident from these plots. Figure 34 !i! i|) shows similar data for high lift, low Mach number conditions° The data selected do not show separated flows on the upper surface as predicted by the AMI CLMAX code (ref. 3), although the angle of attack prediction for the input lift coefficient is good (prior to making lift curve slope corrections). The CLMAX code failed to converge at l_igh angles of attac_for the 7_ thick airfoil. The Squire-Young drag coefficient (_D a v) in CLMAX tended to be optimistic Additional CLMAX cases w_e run to evaluate the predicted maximum lift capability for each tested model. • This code underpredicted the maximum lift coefficient measured in the Ames tunnel by about 10_. (It should be noted that the maximum lift from the Ames ll-foot wind tunnel exceeded that of the UTRC tunnel by 10Y.. ) At a constant lift coefficient the pressures predicted by the NYU transonic (Korn, Garabedian, Bauer) code (ref. 2) are very good, although this code was not formulated for high lift, separated flow conditions and cannot show the same pressure distribution given by the CLMAX code. The TRANSEP code (ref. 1) predicted the pressure distributions well, showing the same or smaller separated zones at the trailing edge than the CLMAX code. The angle of attack correlation would improve if the slotted wall lift curve slope correction was applied to the data.

The surface pressures predicted by the NYU, TRANDES (see ref.

1) and MCMJ-9 GRUMFOIL code (see Eel. 5) correlate very well for the moderate Mach number, moderate lift condition of figure 35. GRUMFOIL provides a better prediction of pitching moment.

Similar correlation exists for the higher Mach number, moderate lift conditions of figure 36. Figure 37 shows the test data - theory comparison for a low lift, high Mach number condition.

The shock position and the pitching moment for the SCI095 airfoil (fig. 37a) is predicted by GRUMFOIL, but GRUMFOIL shows the shock at a more rearward position for the Sc10g4 R8 air- foil. The three codes agree with the test data reasonably well for the transonic airfoils. GRUMFOIL exhibited much better pitching moment correlation than the other codes evaluated.

The NYU, TRANDES and GRUMFOIL predicted the drag divergence Mach number within ±.015. TRANDES tended to underpredict the drag divergence Mach number while the other two programs matched or slightly exceeded the drag divergence Mach number based on test data. The theoretical calculations for the SC1094 R8 airfoil had the largest deviations from the test data. The predicted drag levels for the cases of figures 35-37 were very good.

CONCLUSIONS The test confirmed that the NASA Ames Research Center Eleven- Foot Transonic Wind Tunnel is well suited to airfoil testing.

This test provided data for several airfoil designs including the SSC-AXX and SSC-BXX airfoil families, showing capability greater than that of the baseline SC1095 airfoil in the areas of maximum lift, maximum L/D and drag divergence Mach number.

Several modern airfoil theories were compared with the test data. The AMI CLMAX program had good angle of attack-lift correlation for low Mach number, high lift conditions but underpredicted drag. The Texas A&M TRANSEP program showed good surface pressure correlation, but the cases run failed to give reasonable drag levels. The TRANDES and NYU Transonic codes showed good drag, lift, and surface pressure correlation at low and moderate lifts but failed to predict airfoil pitching moment. GRUMFOIL gives good surface pressure, lift, drag and pitching moment correlation for these conditions.

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r_ r_ _3 eglJ 3_ O _._ r_ I I u_D _3_3 U C -_1 I O C e U w -14 "4"I ',1.4 "H O e U ._1 O .K !i Figure I. TSA installed in the Ames Eleven-Foot Transonlc Wind Tunnel.

.O_IGI_;AL FAGE IS _ QUALm_ 27 It |_

, "it

EX"I'ENOtm - pan _.'U_ leE i.

m_ md, OOR i |

P

.--[ 2O !

J_l_1"_ _4 s4E@l'po_ c-c Figure 2. TSA schematic.

_1 1!

OKIG_ TAL P-:_,"-_..¢ i:; OF POOR _U:kL_Ty

(:]) SC 1095

(_) SSC-A09

(3) ssc-Ao7

(._---

(_) SSC- 808

(E) SC 1094 R8

Figure 3. Airfoil section profiles.

¢ ¢ ® @

Figure 4. Airfoil metric sections.

T-,t.

RN/FT PT = 42 PSF (1.4 ATM) M PT " 30 PSF ( 1.0 ATM) .3 - 2.91 .4 2.07 3.73 .5 3.24 4.47 .6 3.66 5.00 .7 4.00 5.52 .8 4.22 5.80 .9 4.48 - .98 4.51 - 1.07 4.44 - 12- P 1 1.0 ATM .... P • 1.4 ATM 5C1095 I0- SSC -A09 SSC -A07

SSC"908

8.- SC 1094 R8 / _ ' (_)

I'

6. I 4.

i I I I I 0 .2 .4 .6 . e 1.0 !.2 MACH NO.

Figure 5. Test Reynolds numbers.

3O -1 11: OF. POOR QUALITy CALIBRATION FIXTURE AXIAL LOADING NORMAL FORCE LOADING Figure 6. Metric section calibration fixture.

• ._ Ira:: .-', -"; i .

- - a o t_ " STATIC ji .....

b ,_ PRESSUR_i!

' TOTAL PRESSURE PRESSURE t - DEFICIT .JDEFICIT ..... -_, '_ I i.

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STALL CONDITION STALL CONDITION HIGH MACH NUMBER LOW MACH NUMBER Figure 7. Representative manometer board wake rake profiles.

_-= =_--_' ..... -_:-1 _-'--_--

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--- :=-' "--" =---4":--" "-" '- "_-" :---'.__L... _ ..... ._-_ --":a-_--_ "--t ...... 1" -I -" ::= :-"'----; =--L-- (a) Lifg coefficien_ versus angle of atgack Figure 8.-Da_a repea_abiligy - SC1095 airfoil, Mach number z 0._0.

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2"i ^_" _" (b) Drag coefficien_ versus 1±ft coefficient Figure 8.-Continued.

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ORIGINAL PAGE IS OF POOR QUALITY ...... . -. :. -,-- £TT.. "'-- , I

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_--X=: L_-L':i ::E::" _ 5;.4"". -. _:'" (a) r..i. ft Figure 9.-- Balance measurement correlation wi_h pressure measurements.

(hi Drag Figure 9.-Continued.

_:I li OR.iGD_AL PAGE I$ DE J_)OR QUALITy " :'..:_ :;:iV:_.-.--.._ -_-:_:""'.---F.:::--'_I-:::.=:V-::_:F-'_-'.---i!; :-t: i :-::--; - • ' !-::.:"x.-.:.:._.'-i':i ".-'._E-_E]_:- ".-_'llIt] i_i ;":_: ':::t'-'" ::"I"'.-. "::-.1_-:'_:'. i ,::. l , i':,;!: :: :...i.-'" • -- t .-.1.. -I, .... :",i_;..!. ,._.._:.:. -..i. .... :"_ " 1 .... 1 ..... I" " '" "i '" I 1 I ..... ' : . 1" ........... i .................... L-" ............ ! ...........

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_::.7- -- i-.l._ :::-J,-...,:_ ! . .i ... +............ ;.,_l_.-. :: ,. :,: I_;_,,--:-_-; --:_ (c) Pi_ching momen_ Figure 9.-Concluded.

LB

Co'.O08

PT z 30"

1.5-

DRAG

l.O" @ / 0.5- I I I I

.2 .4 .6 .8

MACH NUMBER

Figure 10. Drag for a drag coefficient of 0.008.

(a) Lift coefficient versus angle of attack Figure 11.-- Aerodynamic characteristics of %he 5C1095 airfoil.

--

_ ....... ' .... ....i ..... '::i" " " °..

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.., (b) Drag coefficient versus lift coefficien_ Figure ll.-Con_inued.

4O !I :II (c) Pitching moment coefficient versus lift coefficient Figure 11.-Concluded.

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-1;'. "" "" ::t . "::t "" -._-. ..... _: [- _i--- (aJ Lift coefficient versus angle of attack Figure 12.-Aerodynamic characteristics of the SSC-A08 airfoil.

I II _ORIGrNAL PAGE IS .OF POOR QUALITY (b) Drag coefficient versus lift coefficient Figure 12.-Continued.

!:! :ll ORIGINAL PA,SI_, DE POOR qUALITY _. t_l :+--___- :- .... .- .T.I- -+.',iT"_+-"_"-+'._.,. "+ t;;;.l++".; .:...;. • .... " .... T --.-_-...T _-_:-- _- ° "'x ......

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T :':'1-': """ i -- iii4 - :.g_ ?,.+- :.:';:." . "i.-'-.: .i?i (a_ Lift coefficient versus angle of attack Figure 13.- Aerodynamic characteristics of the SSC-A07 airfoil.

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(b) Drag coefficient versus lift coefficient Figure 13.-Continued.

:I I I: ORIGINAL PAGg I'S DE POOR QUALITY..

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"'::1.. :.;,] 1 - ... _ ..g ,._ :-: "r-'---I (a) Lift coefficient versus angle of air foi i.

Figure 14.-Aerodynamic characteristics of <b) Drag coefficien_ versus lift coefflclent Figure 14 .-Continued.

(c) Pitching moment coefficient versus lift coefficient

Figure 14.-Concluded.

5O !_! I I OF. PC,OR QUALITY (a) Lift coefficient versus angle of attack Figure 15.- Aerodynamic characteristics of the SCI094 R8 airfoil.

! _ / • I ::= __:;

!!1 i t

=--'1 :'- -I E:" 'W" ."=1-- "-'-- :_ ,= I --L- '_""- - _'' : "--- ::i1 ]:2]-:- -,. L.._,_L_ : I ._ _ ;- .--: -,:.-'.:.

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I _.2- ; (b) Drag coefficient versus lift coefficient Figure 15:Continued.

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.:-":_::" F"";:":,: .; ..... ;. ,.

":_['_'_ :: : ':":i-" : .._-" "_" • -_ ::-- L;, 'i " I:-::_i:ii-_-:"T: -':i::'- " ':':: _ - _ _ :"::i.:- --.:]'".i -:-!" :l i!:i!'.'- -:-_._:_ ..--:_F::-_:_:-_:-':_-- : ___-._.

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:=.-; _-. - ___:_T :_,::.__..-..:::l_ : _ ::L:.:_ _ ' .....

t"_: .... _" : "_ • ; l-:.- : " ' t -_--!. ,.-i-:'2:f--_. :,,'::!.... -.-:--: _::i..: ..i: .-.:i:_:i _:.

"':"::- ").': I': -i--:-"':':--- "::'" " (a) Liftcoefficient versus anqle of attack Figure 16.--Aerodynamic characteristics at a Mach number of 0.30.

:1 ! !

t:_ =lE'-_.--- :=-.. .=_=-=_-.:-+_:--.l:=-,:=.f_=:j,. ._ -.,--

.,.,. _.:___

-L_t:_!, !::_ :;LL_--._: L_

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._. .... _:f: I_i_, _ ]., _

=- ..... -t=-'!

(b) Drag coefficient versus Izft coefficient Figure 16.-Continued.

; T.,'7.-._ r . ;7 !_ :.--[]-:id.L .

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• - ! .......

=-:.--- : -._=:.- .-::- Ib-:? :--:-.-_,_{.= l..ot:...

:-,-.;--.._: :---tu (a) Lift coefficient versus angle of attack Figure 17.-- Aerodynamic characteristics at a Mach number of 0.40.

|_.

_..-_-

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• : i,-I- ' ."-- i..'q _,-_ - :_ - i-, - !_ ;-:r:: !24?!_ J --._._:_, ,°..: ._.1 (b) Drag coefficient versus lift coefficient Figure 17.-Continued.

!ii :If '_ "_t."_'F. qUALITY --i: :_: !'::!_ _ li--'-_ b:-:-"_-%-'.: "'!!if:i;! :::i _ i :_: : _" _.;:+-C.. ---r I ,: . _, .:=_-:-L-:-.:.:i.:.. :. .. :.=L .-'::.: : .. "-:': :..._:..:.:..:::., • _:_, _, . _-- :_~ - _:-P=:-_:- -: :..---.:;_::,:=,:-:I::--; ::: ::::'-::_ - _ ......... ,.:-:i=..: ,::: ........... _........ _......... -:F ' i i .i :-"_: -I::-F-L-- !: I'-_ ::-:'¢ :l--:i-l-:.:!:i --.;:-.

,T, t-- ' , S - i .... _* ....

:...._:..:2:_,: !_:,:..:.... _,.::..::.: :: ,._::. ................. ...............

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...... ='L............ -_._.

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(c) Pitching moment coefficient versus l_ft coefficient

Figure 17.-Conclude_.

ORIGINAL PAGE 13 I_ IE_ QuM.,rr'_.

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T.:! '!

• i . = .... | , :.._| .

- V_ l:--I - :'.L": J : ..'L :.= ; l l=."

• ! t'- 'V-'! :- "-- -.'-: _---Y-_ -_L_..-. .t :.-=_- - = !!_:21: -i _+: --"_-- . ;7 , --:::'" ==-"__;_-:.=--i __-" -.:_.-...q--_.

:L= • , - I t - I • -' ._.I .

---_--. ,'" r"Y,. , . t i_!

_-_+.3_! _!_! _!!-_._ _1_.__ :--.I

(a) Lift coefficient versus angle of attack Figure 18.--Aerodynamic characteristics at a Mach n,mmber of 0.50.

ORIGINAE PACE IS OF POOR QUALIT_ Cb) Drag coefficient versus lift coefficie,_t Figure 18 .-Continued, i -"-T'" [:i:.7:: : .... • °..

r_.-..T :-.

• .i--_l

_° | I ° ,

_1 1 -

_ r • i i " ' i : : _- : F" : i- <..- . -. _ 1 ii:_J_i-t: .-I _i_- (c) Pitching moment coefficient versus li_t coefficient Figure IS.-Conc!uded.

"I II ORIGINAL PAGE ;_ _)OR QUAL£%7_ ........................

_.--_-_---q_ _E_- _ _"i-_ .- i

"": =];':!__-::'== .................. _ ..... --- i ,.'_" L - " :::= :":",'-" :,.._:-" -F:.._.=::. i. : ....... :::';' I .............................. _ ..... C-";. h" : :_ (a) Lift coefficient versus angle of attack Figure 19.--Aerodynamic characteristics at a Mach number of 0.60.

(b) Drag coefficient versus lift coefficient Figure 19 _Continued.

"! I ] OF ."'F * T T.._.

POOR Q,_ .... .:" Pitching moment coefficient versus lift coefficient

(c)

Figure 19.-Concluded.

i +.

il coefficient versus angle of attack (aJ Lift Figure 20.-- Aerodynamic characteristics at a Mach number of 0.70.

,_T. _ _. • ,.,--4_f OF POOR ,_'.-,-'*-_-_ _'_.

• _. •G7 -" "...I -- _ L_'..---"- " .:.:-,- [-.- .. --- - z.,. , - _°_.

-, - t-,,-,- .: .._ . ...-._._ .:.-_-: -----:'t""": .... [ ....

.... ,-. Li!i::: _ 7-!:: _ .... " : r-. _ ....

: l ii ...

.... ! ... , u I_ -" : ,..__[:7.

.77 , " .. :i::l i= ;?_" • - _. _ ::U -[4:_7 :: b.."

--- - ;° _ _t22 i: "= ...

._ ;,.--.: :-1 -I--- :-:--:.17-:. :: i_:T-- k= - _--: : .,......,_-.,-

- .... i --

...... '_ "" t"-" ......

N

- t :_._: ._._-- ....... _'-- ,,, .-: _._ .._ _".-_- I_ _"" (b) Drag coefficient versus llft coefficient Figure 20.-Continued.

!!! _| !

o_O_ _'_b _,_.

-'2t.---'_- ...... _" ......... "- .... ---'_---- ....... ":-:: .... --_ :-_:= :-::.- _-{i, ,_/t!ii::'_f_ii'iii::!!i".

Ca) Lift coefficient versus angle of attack Figure 21.--Aerodynamic characterlstics at a Mach number of 0.80.

<b) Drag coefficient versus lift coeffici_n% Figure 2!.-Continued.

7O • _v,_GI?;'.<'_ _-',\GE IS t<K:)R QUA].I'I_ i_&it_!_-:r;-, r=-,- :i:. i: _'_ i- !!-i_i'_ !i!:-:'i:_¸_ ] _ -.-:'.:"_.-.-:_.-..r".:r:--:.'-: ..=.: :,--='-="-:'-.-- -. :i: ':":" - :'i " t • , ......... , ...... I--i. ................ ""i:'" :.:i ::i .--,---E .... -. :, ...... :........ 1...._::! i: : ; '--- -'-_'_- .;:M "_: _-- . . :=-_-_:::_!_!L_i_-_.=_i_!_ _.:_-_÷:4= _:!!_ : : ::-_----:.V,_-"::J2::';-:.::i-;:':_:.E-::GI__ -_,_--_:-:I-_:I: :_:.; :. ' -- _L---_-_'--_-_:z_-4 _',-i:-I_-._-,_ -._-:-_:-_--_:: F::--::._: -' _, - :: t _:=_'_.-_=_±_:'_-::'_.=-_-t-:._--==-_-J_-= :-_::.....;. :_: : t i..... ..i:z] ---,--., -,- [-..--::}:--- -':; ,:" ":."F" ::. .... -_ • (c) . Pitching moment coefficient versus lift coefficient Figure 2i.--Conc!uded.

-- ,_ ... ..2. .

--"P-- "_,--_4" --'-_'_----._%-".-_'--_..:," .i_:o ::.i."_ -_ ;._,.___ ....... z4z_L .... _-: J,::-i .:.b-. :i :: ____ _ _p-_- _ .._,__._ ...... . ...

=-_"_: -': := _-I--': :t:-_............................. _............. ' • * °: o _._ •:::F.:-: " :I :_ _-!:_ ........... -: _i --_i{i:.f:-.-.-_r."-:_ii_f_i!?:! ::;_L: !:iiE:l: ._T., ,: ' : '-- "'" , -... ::,:|: "1 1 :-:_:F:-:: .-:i : { ::_I_.. ::.fi:::-.--.:_:P-F::t,-_.,_l_---- _-:-:-.j-:_:--,-- ! :f:: -_ _::i- ..::i ' t.

.... : ....................................... "- ............... i • " "'" ' I': _ ...... ' ........................ _'_ ..... * ...... _:::-t: ! :--!--::---- -- :=T -]-::_-- :-'=_-:-:_-:_2:;c.--_=|:---.=-" '. " , :-! ....... r.- -::_: --T:P:_] P24i -_ = = _:Z._ff "--_:: !:-=_-:- _:---: ,- .i i : • =--*- .-- _'._. _.-- ;__.,--___. _--._ -.., . .

_ _-_'_$_ _!_ _-_i_i_-_-@c_ '--_--......

Ca) Lift coefficient versus angle of attack Figure 22.-- Aerodynamic characteristics at a Macb number of 0._5.

ORIG_',_'AL PAGE IS OF POOR QUALITY.

---"_ _-_'-_-_:_,_,+__,"+:: +-_::_i : I ::i!_- " _ '

;.-.+. ....... _._++--+++-_ -_'-'::-++-':_"+":.:r' _-: +.. :+:I::+ +++ ++-+: ...+ : -i":--_:-I:-i:++-:_ ::i-': !.:-:.i:::_? i"-'-+:_I!:i:I ?l_:_:I : " t-: I.:_- : .-..l_.i_;l;::-::--,:-_: :i: i:-t_--. !.. :--_+-.--: :-ii.-. l-._iI:: .i.: -_ :.I:- -_ .-, .

_!-+:.T.

":T_ "-:_[:: -t I:: - "-, .', .-': .'r, .- .... i" ,_:':;;.: -.: ; ,':.: :i :::|- .: t_ , --• !- ............ . :-.ii!:_::::}" _:i_ ::_l- !_': ! i , t -=f-..

='+ ;-" : _" -": ":. .... ;: :'_:", i:: i_ _ ;]' ! t.! ;i" " i' .: ""!

• I : i + , --"-'P-" ---- iVlr i :' i.-': _ " -_:+:.:."-:i:.,i!-} • -_ .... ++--.

_ ,+ • . :+:p,,t+- : i.--, "I :':; .- +''+:" '-:'"'+ .....

: _ :_.:.::i::. imP._. :_:j.. :::jil :-:.-!¢::.++ _ -.:_-k: --" --: - -- --;_:-_----i-- :-+'.-----+_ _M-::! :-.- -I. -:- .:+:

?+

(b) Drag coefficient versus lift coefficient Figure 22.-Continued.

-----. _ ++ -q- _.-=__--mm+"__. .........._-_k--iLti?-+.- ++! '_: _=--.-- =.-__-- -:.- ..... ,_.. ....... .-, _,-- _--_--,_,,_--_------i%:: +_-_+-++ I.:+ ++

-= =-_--++--::_=+++-++++ +++i+++++--+--+:+ -#F+.+:.--:.u+ + .+- ++:+ ,:: + ++ ..+.I -+ I

_ _+ _E_+_I__I._ _ _ r. "+ _ :. _ _ I _ ..... _ ' ' "i I ___ ._ + I

+-+++i++ i++1::++: !+++ + .-+_

.,:_.,-.:-+.. ?i ++i_-+l:"--t:-"+++ +++ +++ +.+++ +:.+ +.. + ++

.....+::?i-:+_-:++_+++++;4,+++i_; +

_+,++-l+-:t+:+:l.:+q++i+.++.: t : L_

. '+, I i I _: ::-=.-!: " i:::-_+!tiL _:+:-ii:.:i L :I:._:._t+::-i_P .!!iiii_:,_L:t + ";:_ ; , 1 ++_ " t, j ._. +.. ...... , ,_ ::--_ _-._:: .................

-'-': ..... !I + +=_ _:-:-_.:-:E._- ....................... _: , _++ :.::._.--+._'::.:.p._." ::::_._.:': : .---Tu--l: +-1::-_ :1 :_-.'+_+. ".

_c_ Pitching moment coefficient versus lif_ coeff:clent Figure 22.-Concluded.

.ORIGINAL PAGE DE LK)OR QUALITY ill | ] 2.0 5C1095 SSC-A09

\ ssc-Ao7

1.6 .... SSC-808 _... SC 1094 R8 % 1.2 " "'_.,_. 2_.

- •._---_ -._..

0 .8" ...._._. _- 0.4- I I ! t I i 0 .3 .4 .5 .6 .7 .8 .9 MACH NUMEER Figure 23.-- Variation in maximum lift coefficient versus Mach number.

.028 - 0 SC 1095 [3 SSC -A09 ------- SSC-A07 SSC -B08 • 024- .... SC I094 R8 CZ =0 .020- .016- .012- / B o / 008- .004- 0" I I l I I , .62 .66 .70 .74 .78 .82 .86 .90 M&CH NUME_R Figure 24.-Variation in drag coefficient at zero lift versus Math number.

iz|!II_ SC 1095 SSC -z_09 120- _---- SSC -A07 .... SSC- BOB -""_',_ SC 1094R8 _="'" _.. % I00- "\. _\

\ " \

BO" \"\._.. ' _, _.k \\, MAX LID 60- 40- 20- I I I I I 0 .2 .4 .6 .8 1.0 MACH NUMBER Figure 25.- Maximum L/D versus Mach number.

• 20-

/

• _J SS "_" .08- SC 1095 5

oT_c 8__T TuNe, EL

SC 1094 R8 A I F,,FOI LS .04 - i i I I I .2 .4 .6 .8 I. 0 MACH NUME_ER, M Figure 26.-- Lift curve slope correlation.

_III

I l i t i

MI=ICH O, 40

o = ALF,-3.2

[] = ALF,-O, 2 MACH O. 40

MAOH O, 4.0

" = ALF, 3,0

MAOH O, 40

! o = ALF, 6, l

MACHO. 40

v = ALF, 9,1

!

MACH O. 40

x = ALP, t2,0

+ = ALF, t5,0 MACH O. 40

o5

r, r,1

c3£

n_',.

(a) M = 0.40 Figure 27.--Pressure coefficient distribution for the SCI095 airfc'i.

I I I I I

o = RLF.-3.4 MRCH 0.60

[] = RLF.-O. 3 MRCH 0.60

= RLF. 3. i MROH 0.60

= RLF. 6.2

MRCH 0.60

v = RLF. 9.2 MRCH 0.60

i\i

JJ _

MRCH 0.61

Z x = RLF. 12.0

_J

_f

÷ 0.0 0.2 0.I 0.8 0.6 1.0

X/C

(b) M - O. 60 Figure 27.- Continued.

8O _:I I I @@ i

I I 1

I I

o = RLF.-3.5 MRCH O. 81

= = RLY.-O. 0

MRCH O. 80

" = RLF. 2.2. MRCH 0.81

£N1

[o = RLF. 4.3

MRCH 0.81

v = RLF. 6.3

MRCH 0.81

E--0 Z r,1 I---4 ,II@ -,I,- 0.0 0.2 O.i 0.6 0.8 1.0

X/C

(c_ M = 0.80 Figure 27 .-Concluded.

Sl c_

i i J l I l

o = ALF.-O. 3

MACH O. 40

o = ALF. 3.0 MACH O. 40

MACHO. 40

_. _ _ = ALF 6. I

MRCH O. 40

v - ALF 12

MACHO. 40

MRCH O. 40

x FILF i5

_? ,,

.7

i ÷

o.o 0.2 0.4 XIO o.B o.B i.o

(a_ M = 0.40 Figure 28.-Pressure coefficient distribution for the SSC-A09 airfoil.

q :II

, 1 I I i I

o = RLF.-3.3 MRCH O. 6,0

o = RLF. 0.0 MRCH 0.60

,_ = RLF. 3.3 MRCH 0.60

N

o = flLF. 6.3 MRCH 0.60

v = fiLl. 9.2. MRCH O. 61

x = RLF. i2.0 MflCH O. 61

r.1 \ ,., 0.0 0.2 0.8 0._ 0.6 1.0

X/C

(b) M = 0.60 Figure 28.-Continued.

i I I

I I I

o = ALF.-3.3

MRCH O. 81

: = ALF.-O. i

MRCH O. 81

_ = ALF. 2.2.

MRCH O. 81

d * = BLF 4 3

MACHO. 81

| • •

v = I=ILF. 6.3 M£1CH O. 81

_. R .,_--,

b

I + 0.0 0.2 0.I 0.6 0.8 1.0

X/O

!Ii|i ¸_ Q

i i

I I .... I

o = RL£.-3.3 MRCH O. ¢0

o = RLF.-O. 1 MRCH O. iO

" = RL_ 3.1 MRCH O. 4O

= 8L£. 6.0 MRCH 0.. _0

v = RLF. 9.1 MRCH O. ¢0

x= RLF. 12.0 MRCH O. 40

E,,,4 _

u_& 1 "

C._ I , ib 4, J ._d

0.2 o._ X/C 0.6 0.8 1.o

(a) M = 0.40 Figure 29.-Pressure coefficient distribution for the SSC-A0?

airfoil.

S5 im

o = ALF'.-3.3 MilCH O. 60

a = FILF'.-O. 2. MACH O. 60

=. -_ " = I_LF. 3. I MILCH O. 60

, _ o = FILF. 6.2... MFICH O. 60

!v= FILF. g.2 MFICH 0.60

= FILF. 15.0 MI_gH O. 61

Y

0.0 0.2 0._ 0,8 0.8 1.0

X/C

(b) M = 0.60 Figure 29.-Continued.

, I I I 1

o = RLF.-O. 5 MRCH 0.81

[] = RLF,. 2.2. MRCH 0.81

MRCH 0.81

= " = RLF. 4.2

• o = RLF. 6.'t: MACH O. 81

Z b.,l

,=

Q 4, O.O 0.2 0.{ 0.6 0.8 1.0

X/O

(c_ M = 0.80 Figure 29.-Concluded.

I l I I I

o = RLF.-3.

MROH O. 4,0

= RLF.-O. 2__

MRCH O, 40

• , = RLF. 3.0

MRCH O. 40

o = RLF. 6.1 MRCH O. '_0

v = RL£. 9.1 MRCH O. 4.0

x = RLF. 12.0 MRCH O. 40

+ = RLF. 15.0

I_RCH O. _0

q ÷

o.o 0.2 o._ X/C 0.6 o.e _.o

(a) M = 0.40 Figure 30.-Pressure coefficient distribution for the SSC-B08 88 airfoil.

'Ill

I ! i I 1

o = RLP,-3.4 tlRCH O. 60

= = RLF,-0.2 MRCH 0.60

A = aLP. 3,1 rlRCH 0.60

o= RLF. 6.1 rIRCH 0.60

= RLF. 9.3 NRCH 0,60

x = RLF. 12.0 MRCH O. 61

C3 (:D (b_ M = 0.60 Figure 30.-Continued.

CD !

1 i 1 t

o = ALF.-O.O MRCH 0.81

o= ALF. 2.1 MRCH 0.81

= ALF. 4.3 MROH C.8!

c,i

= ALF. 6.3 MACH 0.81

0.0 ," .2 0.4 on .= r-._. '..F X,/C (c) M = 0.80 Figure 30. Concluded 9O :I 7ll

! I......... 1 t I I

o = ALF,-3,4 MRCH O. 40

o = I=ILF.-O. 2 MROH O. 4.0

"= RLF. 2.9 MRCH O. 40

o = RLF. 6,0 MRCH O. 40

v = RLF. 8.9 MRCH O. 40

x = MLF. 12,0 MROH O. 40

+ = I=ILF, 15.0

MRCH O. 40

t,D !

,-°

r, r,1

o?

I\

n.-" _, Q_ ,

0.2 o.,_ X/C 0.6 o.8 1.o

[a) M = 0.40 Figure 31 .-Pressure coefficient distribution for the SC1094 R8 airfoil 9".

I I I I I

o = ALl=.-3.3

MACH O. 60

= ALF.-O. 3 MRCH O. 60

= ALF. 3.0 MACHO. 60

o = ALF. 6.2 MRCH O. 60

v = ALF. 9. i

MACH O. 60

x = ALF. 12.0

MAOH O. 60

o _. !_ _

0") 4, 0.0 0.2 O.'t 0.6 0.8 1.0

X/O

(b) M = o.60

Figure 31.-Continued.

i:1 |I

o = RLF.-0.3 MACH 0._0

0= ALF. 2.2 MILCH 0.81

_,= RLF. 4.1 MACH 0.81

o= RLF. 6.2 MAOH 0.81

" i 0¢ 4, 0.0 0.2, 0.4 0.6 0.8 t.O

X/C

s f

__l 1 _.L

Du I SC 10_5

o = FILF.-r.6 MFtCH f-. 82

SSC-A0g []= FILF.-0.7 MFtCH O. 82

SSC-Ad7

= FILF. O. 4 MFtCH O. 83

Io= FILF.-O. I MFtCH P. 82

iv = FILF.-C. 8

1 R8 MFtCH O. 83

!=

Z rlt

,.._ M _z'_T,,; ..... m t

.* )

o t

f_ • ,.-, O.C 0._ n.4 F.6 r_._ el.

X/q, [a) M = .825 Figure 32.--Pressure coefficient distribution for low lift at high Mach numbers _I!II @D K !

sc I s o = RLF.-_.6 MRCH 0.85 i I __r..J_ o = RLF.-:.3 NRCH 0.85 I * ssc-_o7 _ = RLF. 0.4 MRCH 0.86 I R , ssc=so8 * = PLF -0 4 MRCH _ _-_ £kl , _ SCl_4RS _ = RLF. O. _ I'IRCH 0.56 i t E--* I , • --r- r,1 . I i .....

I,..-,I i j 1 i t

_ r. _,i_ . i.o. ! .-

i : I m 1' I i .... i "2 ' : O.C 0 " 0.t r._ :'._ ".C ×/C (b) M = .85 Figure 32.-Continued.

aD _L _._____Ira

, i

MRCH O. 85

_C 10g_ o =, ALF.-cj.6

NRCH C. BB

o= FILF.- 1.4

_SC-A0r} MF_Cm O. 9$

= -_ FILF. 0.6

MI_.S_0 aa

o=RLF.-O.5

_-nnn

I MILCH _._7

v= FILF. 0.2

SC1094_ R8

]

I I Z r j

I

ti4 i :'.3 _'_ 1 LL- ! _ !

a oo,I l"*i :I

cn vo_ il

_ g L

Q-.

I I _ T i

• _

i 0.0 0.2 0.I 0,_

X/C

':I I!

[[_ _ _ L 1 1

SC 109_ o =, RLF.--O. 3 M_H O. 93

oo,,-,,_o [] =, RLF. O. 0 M£Ch O. :q._

_C-AO_ !_ = FILF.-O.O MF_CF_ O. 93

ssc-_o_ I_> =, PrI.F_ .-1.4_.

MBCH 0.82

I Z r.l

b

¢..3_

E-

r, 1 0 _ i ()

o o,g

t' 03, :3 _Q O ' _ (> r.=j • O_

I

i 0.0 0,2 0.{ P,_ C,8 i,C

X/C

(d) M = .90 Figure 32. - Continued.

I I i

sc lo9_5 o = PtLF.-t. t MFK]H 0.98

MACH r.38

SSC-A0) _D= RLF.-0.0

M£ ._- 0.99

_SC-A07 A -, 9'JF. --[". _.

4 ssc-_o_ o = RLF.-!.._ M .£-43,H P.98

I _J "..3 5_ ' bJ

_ ! II °

_ , _ ,oo , -!!

I i U-) o_ I! L O_ cL.

:'O I

#_

i

if I

I

CQ t i p,r r ,_ 0-4 P.6

×2C

(,eJ M = . 98 Figure 32.-Continued.

11 fill

L_.L_ L

l J__ _

SC 109J5

o = RLF.-1.2 MROH i. 1

C_J "__AI"I_, ,-1= RLF.-O. 5 MFtCH t. 1.

.' = RLF,-13 4 M,=E]H I. l

SSC-8O8__

o =. RLF.-1.3 M_CH I. i

"T - !

I

Z ...: r

r,_ !

r..l_ r,] .,I,

XIC

(f) M " 1.07 Figure 32.-Concluded.

CD !

I

M= 0.306 RN=3.89x 106 a C_ Cd Cm TEST Q -._5 .086 .0092 -.OI5 .37 .086 .00"/5 -.004 ['_ " KGB z "7 s3 .079 m .002 r,_ TRANSEP -- - • 092 .0061 ..015 CLMAX .oo ( !

I- 0.0 0.2 0.4 0.6 0.8 1.0 X/C (a) SCI095 Figure 33.-Pressure coefficient correlation, M = 0.30, C 1 = O.

i00 _! II CO I M= 0.307 RN= 4.29 x 108 a C I Cd Cm _. TEST C) .01 .142 .0072 -.015 KGB .82 .142 .0083 0 Z '7 1.24 .141 .0065 .009 r,_ TRANSEP -- CLMAX .30 .141 .0069 ..013 C._ r.t..

U') D

a:?

fN t.

0.0 0.2 0._ 0.6 0.8 1.0

X/C

(b) SSC-A09 Figure 33 .-Continued.

I01 co i I I I M= 0.301 RN= 5.51 x 106 a C_, Cd Cm -- TEST 0 -.16 .032 .0086 -.013 KG8 -.06 .032 .0074 -,001 -TRANSEP ---- .30 .043 .0089 .007 CLMAX ---- -.32 .034 .0059 -.009

I I ..... I I

-f 0.0 0.2 0.4 O.B 0.8 l.O

x/c

C3 I i I I ! !

Ms 0,301 RN, 4.75 x 108 a C3. Cd Cm <_ ..13 .001 °0073 .001 TEST I--.- -.22 .001 .0081 .004 KGB .... .74 .007 .0089 .023 TRANSEP !

.... .14 .008 .0084 .004 CLMAX

I I

"9"_ _-_ "_ ;PK_ L--___

_ _ _ _ --_ =; 4- -4- 0.0 0.2 0.4 0.6 0.8 1.0

X/C

(d) ssc-so8

Figure 33. -- Continued.

O I I I I !

Ms 0,303 RN" 4.12 x 108 D TEST O ..22 .021 .0122 ..021 +. KGB -- .11 .021 .0078 ..014 TRANSEP ..... 08 .023 .0111 ..001 CLMAX .... .38 .020 .0o93 -.021

I I l I

:= !

X/C

(e) sclo94 Re Figure 33.- Concluded.

I ]I I ÷

o.o 0.2 o._ X/C 0.6 o.8 z.o

la) scio95

Figure 34.--Pressure coefficient correlation, M = 0.30, C 1 = 1.2.

u; I M= 0.307 RN= 4.29 x 10 b a C_, Cd Cm TEST E) 11.00 1.223 .0131 -.012 KGB t 9.98 1.214 .0116 +.007 _mm cTRANSEP 11.50 1.226 --. .011 LM_ mere 10.40 1.217 .0147 .006 _O i, l !

_OO C._ r, r._ r,1 I r,1 n-" Ct3 0"3 | + 0.0 0.2 0.8 1.0

o._ X/C 0.6

(b) sSC-AO9

Figure 34.-- Continued.

I I, X 108 M= 0.301 RN = 5.51 G C d Cm

c_

TEST 0 11.00 1.149 .0177 -.015 KGB TRANSEP ---- 10.18 1,151 -- ..007--_ CLMAX -- 8,70 .975 .0100 .019

Io I

E-4CD C._ r.s..

r._ I r_, CF) i.,iI ° t% I \ c>,._ • I | • c; I C) "N

o.o 0.2 o._ X/C o.s 0.8

_c_ SSC-A07 Figure 34.-Continued.

i07 u'_ I M= 0.303 RN= 4.75 x 106 o C_ Cd Cm

Ill

TEST 0 13.06 1.205 .0178 .013 KGB 10.56 1.193 .0120 .024 i TRANSEP ------ 12.56 1.206 -- .039 i CLMAX ''-- 12.20 1.193 .0157 .033 -- !

L b..

W C.._ i r._ o:: 0") o') e,ql ¸ ,q.

Q q,o -t 0.0 0.2 1.0

o._ ×/O o.e

(d) SSC-BO8

Figure 34.- Continued.

i1 | ]_ , 1 i q14 I .

! I M = 0.298 RN= 4.12 x 106 (Z C_. Cd C m 0 12.09 1.261 .0154 -.001 TEST

il

-- 10.52 1.254 .0102 .006 KGB TRANSEP .... 10.72 1.254 -- .012__ F_ CLMAX -----; 11.10 1.287 .0115 .023 I ,I l l-t _ !I I if-)

%

a_ I J __. --_, II ÷ 0.0 0.2 0.8 t.O

o._ ×/C o.e

(e} SC1094 R8 Figure 34.- Continued.

¢= ' 'L _o M • 0.2S8 RN • 4.12 x 104

i 1

a (:2. ¢d ¢m- I TEST !i KGE t 1.811 -- ._30 l&O0 1--=04 .0"I_.$ ._Z_ ¢LMAX

I

eel I Z n_ =.

U3 _3 I .._..- - -.__..._ _ i--cr-

_iff

N g _'o_

_'.= o'._ ×/00.e

(f) sol094 R8, cI = 1.5 Figure 34.-Concluded."

ii0 I 111 0D !

M= .401 RN= 3.63 x 106 a C._, Cd Cm 0O TEST O 6.13 .745 .0080 -.012 KGB 5.67 .742 .0096 .005-- I TRANDES .... 5.70 .744 .0091 ..004 GRUMFOIL ---,'_ 5.28 .741 . )092 :.012 I C= (%1 Z r,1 O r.

r, ,...

r,1 U') r.1 rv" 0..

,q, 4- 4.

0,0 0,2 O.i 0.6 0.8 i.0

×/c

(a) SCI095 Figure 35.-Pressure coefficient correlation, M = 0.4, C 1 = .7.

iii r.D M= .399 RN: 3.85 x 106 o Ci _ Cd C m CO TEST E) 6.23 .785 .0066 -.016 ------- 6.04 .781 .0099 .004- KGB I TRANDES -- -- -- 6.20 .785 ,0094 -.008 GRUMFOIL 5.55 .779 .0089 -.012 O p._ Z L_1

\

I,,-q r, r, C_e I r,1 U') r,1 ,i + P,0 t- 0.0 0.2 {:}.4 0.8 1.0

(b) SSC-A09

Figure 35.-- Continued.

I 11 I I I | !

RN=4.83x 106 M = .406 C m _ Cd a Cj_ ._ -.015

0 6.03 .875

TEST .00_ 5.12 .673 .OOO_ _KGB ._ -.011 5.4O .674 TRANDES .0081 -.008 4.88 .870 GRUMFOIL _- ---

I I

I I,,

I

,.: f

4- 0.0 0.2 0.'_ 0.6 0.8 1.0

×/C

(c} SSC-A07 Figure 35.- Continued.

ll3 M= .395 RN = 4.30 x 106 _D Cm m o C_ Cd I TEST O 6.08 .614 .0101 .001 KGB 5.01 .611 .0092 .012 m TRANSEP "- "- " 5.50 .613 .0085 .014 CLMAX 4.84 .607 .0082 .007

\

i 0.0 0,2 O.'i 0.6 0.8 l.O

X/C

(d) SSC-B08 Figure 35.- continued.

_I li _D l M = .402 RN = 3.62 x 106 a C f. Cd Cm 0D TEST Q 5.97 .879 .0079 -.015 KGB - 5.25 .676 .0100 ..003 --- I TRANDES .... 5.00 .888 .0069 -.007 GRUMFOIL ---.--- 4.86 .675 .0091 -.019 A I !

E_ a i I r,1 j

- I

k

%

-I

( r,1 rw _J ¢N ira4 4.

0.0 0.2 O.i 0.6 0.8 1.0

x/c

(e_ SCI094 R8 Figure 35.-Concluded.

_D r'3 i RN=4.90x 106 M = .601 o C_, Cd Cm CD 3,14 .498 .00cj4 ..016 TEST Q 3.08 .497 .0087 .009 - KGB I 3.00 .497 .0082 ..003 TRANDES ....

2.67 .494 .0083 -.013 GRUMFOIL ----- I C3 Z ' t,--,-I CD r, r, r_ \ r,1 L ) n-- 0"3

r,,?

n,- Q..

L) (-_ l w,,4 + 0.0 0.2 0.4 0.6 0.8 1.0

X/C

(a) 5C1095 Figure 36.--Pressure coefficient correlation, M = 0.6, C 1 0O e I i I RN= 5.16 x 10 6 ' M : .599 Q

c£ cd c m

3.26 .519 .0077 -.020 TEST O O 3.27 .517 .0092 .005 " KGB .519 .0083 -.005 TRANDES 3.30 I 2.76 .513 .0079 -.013 GRUMFOIL o ,t, .!..0 0.6 0.8 0.0 O.2 0._

X/C

(b) SSC-A09 Figure 36. - Continued.

Q r4 I I I ! I M = .601 RN : 6.45 x 106 a C t Cd Cm - TEST O 3.15 .348 .0072 -.015 c: KGB 2.09 .347 .0076 .000 TRA;JDES "--'"" 2.00 .346 .0071 -.009-- I GRUMFOIL '-'"-" 1.82 .343 .0071 -.009

I I 1

Z r,1 im,,I i-,-t r, I r.,_ C_ i ,t- 0.0 0,2 0._ 0.6 0.8 1,0

×/C

I18 i| | il I [ i

I

M = .608 RN= 5.95 x 106 a" CjL Cd Cm- TEST Q 3.07 .358 .0079 .002 (:: KGB "--'-- 2.29 .354 .0081 .010__ r_ " I"RANDES ....

2.70 .354 .0075 .019 ' GRUMFOIL ------- 2.21 .348 .0074 ,007 [.-., r._

i

l--q 'l ' _ '-- I I • , [ 0.0 0.2 0.I 0.6 0.8 1.0

X/C

(d) SSC-B08 Figure 36.- Continued.

I19 0O I i I I i i i M = .604 RN = 4.84 x 106 o C./. Cd C m _ m TEST O 3.03 .403 .0081 -.019 KGB -.----- 2.54 .402 .0088 -.006 C: TRANDES ....

2.19 .411 .0069 -.016 _ '--GRUMFOIL - mmwnm romp 2.13 .399 .0080 -.023 I

I I

.....

Z r,1

A

I ( C3

° J

cu _.

f3_ O ¸ ::::3 ' 0") OD r,1 n,- t'_ ÷ 0.0 0.2 0._ 0.8 0.8 1.0

X/C

(el scio94 R8 Figuz'e 36. -- Concluded.

::1 I i M s .825 RN = 5.71 x 108 a CIZ Cd C m TEST -.65 ..043 .0131 -.024 KG8 -.64 ..043 .0127 -.014 t TRANDES ....

-.90 ..057 -.009 GRUMFOIL -""- -.73 ..i)46 .0155 -.024 Z r,.l h.1 m

= ]/

÷ ÷ 0.0 0.2 0.'_ 0.6 0.8 I .O

×/C

( a_ SCI095 Figure 37.-- Pressure coefficient correlation, M = 0.825, C 1 = O.

i_!

0D I I I I I M: .828 RN= 6.08 x 106 o C_ Cd Cm TEST C -.84 .052 .0083 -.02S O KGB ..04 .052 .0096 .002 "TRANDES ..... .50 .046 -.005 GRUMFOI L "'-" -.70 .046 .0080 -.021 [._

I I 1 I

Z r.D (_ I--I r, r, r, 1 CD D uOo 0-) I r,l

,/

('_ B--I 4" 0.0 0.2 0._ 0.6 0.8 1.0

X/C

(b) SSC-A09 Figure 37.--Continued.

TEST KG8 0.0 (c_ Ssc-AO7 FSgur e 37.--c°ntinued" ]=23 ¢D I M:.824 RN: 6.87x 106 a C_ Cd Cm C_ TEST O -.12 -.012 .0080 .002_ KGB -.20 -.012 .0080 .014 I TRANDES .... .30 -.014 .019 GRUMFOIL -.50 -.018 .0081 .000 Z

h

0.0 0.2 0.t 0.6 0.8 1.0

X/C

(d) SSC-BO8 Figure 3?.-Continued.

'I1i ¢D 4l M: .827 RN: 5.87 x 106 i a C_ Cd Cm ¢D TEST Q -.78 ..180 .0217 -._-- KGB -- -.95 -.177 .0209 -.023 I _ANDES ..... 1.51 -.195 -.023 GRUMFOIL -..-.,.,. -.70 -.182 .0261 -.031 Z r,1 ii=t ii-i 4.

0.0 0,2 O.i 0.6 0.8 1.0

×/C

(e) SCZ094 R8 Figure 37.-Concluded.

APPENDIX A TABULATED DATA _'I I I- Headin@ Description for Tabulated Data ALPHA Angle of attack, deg CDBAL Balance - derived drag coefficient CDP Wake rake - derived drag coefficient CLBAL Balance - derived lift coefficient CLP Airfoil surface pressure - derived CMBAL Balance - derived quarter chord pitching moment coefficient CMP Airfoil surface pressure - derived pitching moment coefficient Configuration I = SCI095 Configuration 2 = SSC-A09 Configuration 3 = SSC-A07 Configuration 4 = SSC-B08 Configuration 5 = SCI095 R8 Configuration 6-10 = SSC-A09 Out-of-Contour Test Configuration (See page 9 and Table IV) L/D BAL Balance - derived lift-drag ratio L/D P Surface and wake rake pressure derived lift-drag ratio _CH Free stream Mach number PT Data point number within each run RN Reynolds number based on airfoil chord RUN Test run number (see also Table V) IODOO0 _OOa_O_OO_N_ _O0 DO_OOO0600OO_ I° I" I' t' l' I" I" O" I" l' I' a* l'O 0 o' I' I" l" I' I* I* I' |lllOO_li_OOOO0| O0||OOO_I|O|III§ i OOO_O_OO_O0_OOO0 OO0_OO_O00000OOm I'I°I°I'I'I'I'OQJI°I'I'I'I'I" _l'l'l'l*l'/OOOl'l'l'l','l'l" ! ,u,O,iJt(J, _ _ _ 0 _o_ .......

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REFERENCES I) Carlson, L. A.: A FORTRAN Program for Transonic Airfoil Analysis Or Design. NASA CR-2821, June, 1977.

2) Bauer, F; Garabedian, P.; Korn, D. and Jamison, A.: Super- critical Wing Section II, Lecture Notes in Economics and Mathematical Systems. Vol. 108. Springer-Verlag, 1975.

3) Maskew, B.: CLMAX Program Description.

AMI Report 7711, December, 1977.

4) Hicks, R. M; and Vanderplaats, G. N: Application of Numer- ical Optimization to Design of Low-Speed Airfoils. NASA TM X-3213, 1975.

5) Melnick, R. E; Chow, R. R; Mead, H. R.; and Jameson, A.: An Improved Viscid/Inviscid Interaction Procedure for Transonic Flow Over Airfoils. Grumman Aerospace Corporation, February, 1980.

6) Jepson, W. D.: Two Dimensional Test of Four Airfoil Config- urations With An Aspect Ratio of 7.5 and a 16-inch Chord Up to a Mach Number of i.I. SER-50977, Final Report for Con- tract N60921-73-C-0057, April 5, 1977.

7) Allen, H. J: and Vincenti, W. G: Wall Interference In A Two-Dimensional-Flow Wind Tunnel, with Consideration of the Effect of Compressibility. NACA Report No. 782, 1944.

8) Hilton, W. F: High Speed Aerodynamics.

Longmans, Green and Co., 1951.

9) Bazin,M.: A Critique of Transonic Airfoll Testing Tech- niques. Part I, System of Industrial Tests in S3MA.

L'Aeronautique et L'Astronautique, No. 31, pp. 1-8, Voi. 7.

1971.

_i l_ 2. Governnwmt _ No. 3. R_t's Cemlog No.

1. Report No.

NASA CR-166587 §. Rmm't Dete Evaluation of Advanced Rotorcraft An Experimental February 1 984 6. IN._ormir_ Orgo_J.ion Code Airfoils in the NASA Ames Eleven-Foot Transonic Wind Tunnel 8. PIrforminli OrgJnization R_ort No.

7. Author{s) SER-SI Ol 06 Robert J. Flemming 10. Work Unit No.

T3334Y 9. l_rfQtm_rtg Ofgm_izltion Name and Addreu Sikorsky Aircraft Division 11. ContrK't or Grlnt No.

United Technologies Corporation 14800-039 N. Main St., Stratford, CT 06602 13. TyI_ of Report and P_tiod Covered MCOnt a epq t 12. S_mmi_ Ageflcv Name and Addrm

arc B - orIT 19 3

National Aeronautics and 14. Sponsorit_l _ Code Space Administration Washington, D.C. 20546 15. _D_ementary NoI_ Point of Contact: Raymond Hicks, Applied Aerodynamics Branch NASA Ames Research Center, M/S 227-6 965-5656 Moffett Field_ CA 94035 1415) 16. Abstract Five full scale rotorcraft airfoils were tested in March and April 1982 in the NASA Ames Eleven-Foot Transonic Wind Tunnel for full scale Reynolds numbers at Mach numbers from 0.3 to 1.07. The models, which spanned the tunnel from floor to ceiling, included two modern baseline airfoils, the SC1095 and SC1094 R8, which have been previously tested in other facilities.

Three advanced transonic airfoils, designated the SSC-A09, SSC-A07, and SSC-B08, were tested to confirm predicted performance and provide confirma- tion of advanced airfoil design methods.

This test has shown that the eleven-foot tunnel is suited to two-dimensional airfoil testing.

18. Ois_i_tim $_lement 17. Key W_ (Suggmt_ _ Author(s)) Airfoils Wind Tunnel Test Unclassified - Unlimited Correlation Aerodynamics Helicopters Subject category 02 Transonic Airfoils 22. Price" 19. S_'_riw Olmif. (d thi. r.Dort] _. Securiw Claret. iof thi* I_11) 21. NO. of Pagm Unclassified Unclassified "For sale by the Nltionlt Technicll Information Service. Springfield, Virginil 22161 ::1 l i-

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