Large-scale Advanced Prop-fan (LAP) high speed wind tunnel test report

CONTENTS

PAGE

SECTION

ABSTRACT 1 l.O 2.0 SUMMARY INTRODUCTION 3.0 4.0 PROP-FAN DESCRIPTION 4.l General Description 4.2 SR-7L Blade 4.3 Pitch Change Actuator 4.4 Pltch Control 18 4.5 Hub and Blade Retention 18 4.6 Spinner INSTRUMENTATION DESCRIPTION 5.0 5.1 General Description 23 (DAS) 5.1.1 Electronlc Data Acqulsltlon System 5.1.2 Prop-Fan Diagnostic Monitoring Instrumentation 5.2 Steady Pressure Measurement System 5.3 Unsteady Pressure Measurement System TEST FACILITY DESCRIPTION 6.0 6.1 Wind Tunnel Description 31 6.2 Test Rig Description 6.3 Test Rig Modifications SPINNER AND CENTERBODY DRAG MEASUREMENT 7.0 7.1 Test Rig Vlbration Survey 7.2 Wind Tunnel Corrections 7.3 Test Rig Corrections 7.3.1 Test Objectives ?.3.2 Test Procedure 7.3.3 D1scusslon and Results PRECEDING PAGE BLANK NOT VILMED iii

CONTENTS (Continued)

SECTION PAGE

8.0 BLADE STRUCTURAL DYNAMIC EVALUATION 71 8.1 Test Objectives 8.2 Test Procedure 8.3 D1scusslon and Results 8.3.1 General Discusslon 8.3.2 Data Reductlon 9.0 AERODYNAMIC PERFORMANCE EVALUATION I05 9.1 Test Objective m_ I05 9.2 Test Procedure I05 9.3 D1scusslon and Results 106 9.3.1 Data Reductlon Procedure 106 9.3.2 Data Presentation 10.0 BLADE SURFACE STEADY PRESSURE MEASUREMENT 10.1 Test Objective 10.2 Test Procedure I0.3 D1scusslon and Results 11.0 BLADE SURFACE UNSTEADY PRESSURE MEASUREMENT 157 11.1 Test ObJectlves 11.2 Test Procedure II.3 Discusslon and Results 12.0 CONCLUSIONS 12.1 Blade Structural Dynamic Evaluatlon 12.2 Aerodynamic Performance Evaluatlon 12.3 Blade Surface Steady Pressure Measurement 12.4 Blade Surface Unsteady Pressure Measurement List of Symbols References APPENDIX A Chronologlcal History of Test iv

ILLUST TIONS

PAGE

FIGURE

3-I 6 Prop-Fan Concept 3-2 LAP Program Elements 3-3 Prop-Fan Test Assessment (PTA) Program Elements 4-I 12 Large-Scale Advanced Prop-Fan 4-2 Design Requirements and Goals Summary 4-3 Features of the SR-7L Blade Construction 4-4 LAP Pitch Change Actuator 4-5 Motor Driven Pitch Control Mechanism 4-6 LAP Control Schematic 19 4-7 SR-7L Hub 4-8 SR-7L Spinner 5-I H.S.D. and O.N.E.R.A. Instrumentation List 5-2 LAP Instrumentation System Schematic 25 5-3 LAP Blade Pitch Angle Measurement System 5-4 Steady Pressure Measurement System 6-I Modane-Avrleux Aerothermodynamlc Test Center 6-2 SI-MA Large Atmospheric Wind Tunnel 33 6-3 SI-MA Wind Tunnel Schematic 34 6-4 SI-MA Mind Tunnel Nozzle Inlet to Test Section 6-5 Tunnel Driving Power and Reynolds Number as a Function of Math number 36 6-6 37 Prop-Fan Installation In Test Sectlon/Chariot No. 3 6-7 Prop-Fan Drive System Engines and Gearbox 6-8 Prop-Fan Test Rig Power Llmitations 39 6-9 Prop-Fan Drive System 41 6-10 Prop-Fan Test Rig Dr|ve Train "Balance" Schematic 6-11 Origlnal Test Rig Configuration 43 6-12 Modlfled Test Rig Configuration 7-1 Axial Mach number Distribution for Freestream Mach number .354 7-2 Axial Mach number Distribution for Freestream Mach number .615 7-3 Axial Mach number Dlstributlon for Freestream Math number .721 7-4 Axial Mach number Distribution for Freestream Mach number .754 7-5 Axial Mach number Dlstrlbutlon for Freestream Mach number .788 52 7-6 Ax|al Mach number Dlstrlbutlon for Freestream Mach number .829 7-7 Splnner/Body Tare Test 56 7-8 Test Rlg Support Structure for Splnner/Body Tare Test 7-9 Prop-Fan Blade Stubs V ¸

ILLUSTRATIONS (Continued)

FIGURE PAGE

7-10 Centerbody Surface Static Pressure Tap Locations (Forward Section) 59 7-11 Centerbody Surface Static Pressure Tap Locations (Mid and Aft Sections) 60 7-12 Spinner Bulkhead Static Pressure Tap Locations 7-13 Prop-Fan Test Rig Forces 7-14 Spinner Drag Coefficient vs. Corrected Mach number 7-15 Centerbody Drag Coefficient vs. Corrected Mach number 8-1 Test Rig Support Structure for Structural Dynamic Evaluatlon (2 and 4 Blade Configurations) 72 8-2 Test Rig Support Structure for Structural Dynamic Evaluation (8 Blade Configuration) 73 L 8-3 Two Blade Test Configuration 8-4 Four Blade Test Configuration 8-5 Eight Blade Test Configuration 8-6 Strain Gage Arrangement for Two Blade Configuration 8-7 Strain Gage Arrangement for Four Blade Configuration 8-8 Strain Gage Arrangement for Eight Blade Configuration 8-9 Test Points and Blade Shank Moments (2 Blade Configuration) 81 8-10 Test Points and Blade Shank Moments (4 Blade Configuration) 8-11 Test Points and Blade Shank Moments (8 Blade Configuration) 8-12 Comparison of Signal to Noise for Flatwise Shank Moment Gages (0 ° Inflow Angle) 85 8-13 Trends of IP Response for 3° Angular Inflow (Shank Moment) 8-14 Trends of IP Response for 3° Angular Inflow (Radial Bending) 95 8-15 Radial Bending vs. Span for 3° Angular Inflow 8-16 Comparison of SR-7L Natural Frequency Test Results to Predictions 8-17 Test vs. Analysis (Radial Bending vs. Span) I00 8-18 Test vs. Analysis (Effect of Power) I01 8-19 Test vs. Analysls (Effect of Mach No.)

9-1 LAP Blade Angle Potentlometer Output I07 9-2 Cp and CTNET VS, 3, 4 Blades, MN = .2, _ = 0 = Ii0 9-3 Cp and CTNET VS, J, 4 Blades, MN = .5, _ = 0 ° Iii 9-4 Cp and CTNeT VS. _, 4 Blades, MN - .8, _ - 0 ° 9-5 C, and CTNET VS. J, 2 Blades, MN - .5, _ = 3 ° 9-6 CTNET VS. Cp, Constant O, 4 Blades, MN = .2 9-7 CTNE, VS. C,, Constant O, 4 Blades, MN - .5 vl

ILLUSTRATIONS (Continued)

FIGURE PAGE

9-8 116 CTNET VS. Cp, Constant J, 4 Blades, MN = .8 9-9 CTNET VS. Cp, Constant J, 2 Blades, MN - .5, 9 = 3° 9-I0 CTNET VS. Cp, Constant J, 8 Blades, _ = 0 ° 9-ll CTNET VS. J, Constant Cp, 4 Blades, 9 = 0° 9-12 CTNET VS. J, Constant Cp, 8 Blades, _ = 0° 9-13 Comparison of Measured and Predicted Performance, 4 Blades 9-14 Compar|son of Measured and Predicted Performance, 8 Blades lO-I Steady Pressure Test Rig Support Structure 128 I0-2 Prop-Fan Steady Pressure Test Set-Up 127 I0-3 LAP Blade Installation, Steady Pressure Test (2 Blade) ]0-4 LAP Steady Pressure Blade (Camber Side) 10-5 LAP Steady Pressure Blade (Face Side) 131 10-6 Steady Pressure Blade (Camber Side) 10-7 Steady Pressure Blade (Face Side) ]0-8 Arrangement of the Pressure Taps on the Blade Surface-Projected View 134 I0-9 LAP SR-TL Steady Pressure Distribution of On-Llne Data (Operating Condition No. 7) lO-lO LAP SR-TL Steady Pressure Distribution of On-Llne Data (Operating Condition No. 2) lO-ll LAP SR-7L Steady Pressure Distribution of On-Llne Data (Operating Condition No. 3) 142 lO-12 LAP SR-7L Steady Pressure Distribution of On-Llne Data (Operating Condition No. 4) lO-13 LAP SR-TL Steady Pressure Distribution of On-Line Data (Operating Condition No. 5) lO-14 LAP SR-7L Steady Pressure Distribution of On-Line Data (Operating Condition No. 6) lO-15 LAP SR-7L Steady Pressure Distribution of On-Line Data (Operating Condition No. 9) lO-16 LAP SR-?L Steady Pressure Distribution of On-Line Data (Operatlng Condition No. 8) lO-17 LAP SR-7L Steady Pressure Distribution of On-Line Data (Operating Condition No. 7) lO-IB LAP SR-7L Steady Pressure Distrlbutlon of On-Line Data (Operating Condition No. lO) ]0-19 LAP SR-TL Steady Pressure Distribution of On-Line Data (Operating Condltlon No. 11) 150 I0-20 LAP SR-7L Steady Pressure Distribution of On-Line Data (Operating Condition No. 13) vii

ILLUSTRATIONS (Continued)

FIGURE

PAGE

10-21 LAP SR-7L Steady Pressure Distribution of On-Line Data (Operating Condition No. 12) 152 I0-22 Repeatability Problem for Static Low Power Point 153 10-23 Repeatability Problem for Static High Power Point 154 10-24 Illustration of Repeatability at My = .20 155 II-I Unsteady Pressure Test Rig Support Structure 158 !I-2 Prop-Fan Unsteady Pressure Test Set-Up 11-3 LAP Blade Installation Unsteady Pressure Test (2 Blade) 11-4 LAP Unsteady Pressure Blade (Camber Side) II-5 LAP Unsteady Pressure Blade (Face Side) 164 11-6 Unsteady Pressure Blade, Transducer Locations (Face Side) I!-7 Unsteady Pressure Blade, Transducer Locations (Camber Side) 11-8 LAPISR-TL Unsteady Pressure Blade Transducer Mounting 11-9 Unsteady Pressure Test Set-Up With Wake Generator 11-10 Examples of Measured Unsteady Pressures 11-11 Illustration of the Terms "Advancing" and "Retreating" for Angular Inflow

TABLES

PAGE

TABLE

7-I Splnner/Centerbody Drag Test Points vs. Corrected Mach No.

8-I Operating Conditions for Mn >.l and Sample IRP Tabulations (2 Blade Configuration) 8-II Operating Conditions for Mn >.l and Sample IRP Tabulations (4 Blade Configuration) 8-III Operating Conditions for Mn >.l and Sample IRP Tabulations (8 Blade Configuration) 8-IV Measured Vibratory Peaks vs. Limits P-Order Content of Strain Gage Output at 3° Inflow Angle 8-V (2 Blade Configuration) 8-Vl Comparison of IP Content of Strain Gage Output Obtained by Two Methods 8-Vll Comparison of Test vs. Analysis (IP) I0-I Operating Conditions for Blade Steady Pressure Testing (2 Blade LAP Propeller) II-I Operating Conditions for Blade Unsteady Pressure Testing (2 Blade LAP Propeller) ll-ll Chart of Transducers from wh|ch Data was Acquired vs.

Test Condition ix/x l.O ABSTRACT High Speed Wind Tunnel Testing of the SR-7L Large Scale Advanced Prop-Fan (LAP) is hereln reported. The LAP |s a 2.74 meter (9.0 FT) diameter, 8-bladed tractor type rated for 4475 KW (6,000 SHP) at 1698 RPM. It was designed and built by Hamllton Standard under contract to the NASA Lewis Research Center. The LAP employs thin swept blades to provide efficlent propulsion at flight speeds up to Mach .85.

Testing was conducted in the ONERA SI-MA Atmospheric Wind Tunnel in Modane, France. The test objectives were to confirm the LAP was free from high speed classical flutter, determlne the structural and aerodynamic response to angular Inf|ow, measure blade surface pressures (static and dynamic) and evaluate the aerodynamic performance at various blade angles, rotational speeds and flight Mach numbers.

The measured structural and aerodynamic performance of the LAP correlated well with analytical predictions thereby providing confidence in the computer codes used for design. There were no signs of classlcal flutter throughout all phases of the test up to and including the 0.84 maximum Mach number achieved. Steady and unsteady blade surface pressures were successfully measured for a wide range of Mach numbers, Inflow angles, rotatlonal speeds and blade angles.

No barriers were discovered that would prevent proceedlng with the PTA (Prop-Fan Test Assessment) Flight Test Program scheduled for early 1987.

1/'2 Z 2.0 SUMMARY This report describes the procedures followed and results obtained during High Speed Wind Tunnel Testing of the SR-7L Large Scale Advanced Prop-Fan (LAP). The LAP Is a 2.74 meter (9 foot) diameter, 8 bladed advanced propeller designed to attain high propulsive efficiency at flight speeds up to Mach .85. The Prop-Fan achieves this superior speed and efficiency by employing thin swept blades and high disc loading. The High Speed Wlnd Tunnel Test was conducted in the ONERA SI-MA Large Atmospheric Wind Tunnel facility In Modane, France.

Testing was accomplished during two separate tunnel entries. This was necessitated by a test rlg failure during the first entry. Prior to Interruption of the test in early 1986,.a11 structural dynamic, aerodynamic performance and a limlted amount of blade steady pressure testing was completed. Testing resumed in early 1987, and culminated in the completion of the steady and unsteady blade pressure tests. A complete chronological history of both tunnel entries Is provided in Appendix A.

The purpose of the Wind Tunnel Test was to confirm the LAP was free from high speed classical flutter, determine the structural and aerodynamic response to angular inflow, measure blade surface pressures (both steady and unsteady) and evaluate the aerodynamic performance at various blade angles, rotational speeds and Mach numbers. Results from these tests would assist in determining the readiness of the LAP and Its Instrumentation systems for follow-on Prop-Fan Test Assessment (PTA) flight testing. The Wind Tunnel Test was accomplished in two phases. In the first phase, structural dynamic and aerodynamic performance data were collected concurrently, for the 2, 4 and 8 blade configurations, over a wide range of blade angles, Mach numbers and rotational speeds. Due to rig constraints, structural dynamic evaluatlon of the LAP operating at a fixed angle of attack was limited to the 2 bladed Prop-Fan configuration at a 3° inflow angle. During the second phase of testing, a specially fabricated static pressure tapped blade was employed to map the blade surface steady pressure distributlon for a range of operating conditions. Due to drive system power limitations, testing was accomplished utilizing a two bladed Prop-Fan configuration to provide blade loadings simulating the take-off, cutback and design cruise conditions. The final phase of testing, also utilizing the two blade configuration for reasons mentioned above, employed another specially instrumented blade incorporating high frequency response pressure transducers. Data from these transducers was used to define and evaluate the blade surface unsteady pressure distribution for the same operating conditions run during the steady pressure test.

Unsteady pressure testing included evaluating the effects of a wake In the propeller inflow at a 3° angle of attack.

2.0 (Continued) Results from the High Speed Wind Tunnel Test demonstrated that the SR-7L Prop-Fan was free of high speed blade flutter over the entire operating envelope tested. Additionally, all measured blade surface and blade shank strains were well below allowables set prior to testing. Good correlation was found between measured and analytlcally predicted IP strain sensitivities for the SR-7L blade. Results confirmed that IP strain peaks Inboard on the blade and lessens near the tip, and that, in general, blade strains were found to increase with power and Mach number. Measured aerodynamic performance for the four blade configuration corresponded well with analytlcal predictions over the entire range of points tested. Though good agreement for the eight blade configuration was found at Mach numbers of .70 and .73, performance was slightly underpredlcted at .5 Mach number. Steady and unsteady blade surface pressure measurements were successfully collected for a wlde range of Mach numbers, rotational speeds, blade angles and inflow angles. In addition to confirming the presence of tip edge and leadlng edge vortex flows at Mach numbers of zero and .2, shock waves were evident at the traiIlng edge at .7 and .78 Mach number. Unsteady pressure responses were clearly present as evidenced by a dominant l-P response in the angular inflow data and significant 2-P response in the wake inflow data. Sinusoldal response was evident on the pressure (face) side of the blade for all cases examined for angular inflow conditions and on the suction (camber) side under low loading conditions. Under high loading condit|ons, the suction (camber) side exhibited non-slnusoldal response resulting from the presence of tip and leading edge vortices.

3.0 INTRODUCTION The Prop-Fan, a high speed, high efficiency alrcraft propulsion concept was launched during the "O|l Crunch" days of the mid 19?O's. In response to the national need to reduce fuel consumption, Congress directed NASA to address a series of aircraft related technologies aimed at increasing the fuel efficiency of airline operation. In response, NASA created the Aircraft Energy Efficiency (ACEE) program which addressed fuel savings through advancements in both airframe and engine technology. The element of the ACEE program offering the greatest potential fuel savings was the Advanced Turboprop Program (ATP) as described In Reference I. The NASA Lewls Research Center had total responsibility for the ATP project which is summarized in Reference 2. The objective of the ATP was to demonstrate technology readiness for efficient, reliable and acceptable operation of advanced turboprop-powered commercial transports at cruise speeds up to Mach 0.8 and at altitudes above 9,BOO meters (30,000 ft.) while maintaining cabin comfort levels (noise and vibration) comparable to those of modern turbofan-powered aircraft. The technology would also apply to possible new military aircraft for a variety of missions. Out of this project evolved the Prop-Fan concept.

Although high propulsive efficiency from turboprops was nothing new, the standards of high cruise speed and cabin comfort set by the contemporary turbofan powered alrcraft were beyond the capability of any turboprop powered aircraft. The Prop-Fan concept evolved to satisfy the requirements of high speed and altitude with improved efficiency while maintaining a high degree of cabin comfort. It is characterized by the large number of blades (8 or lO), thin airfoil sections, and swept blade planforms (Figure 3-I).

Once the concept and its benefits were defined, NASA conducted a systematic program to verify that the predicted benefits could be achieved and that there were no unsolvable problems in Implementing the concept. The potential benefits of the Prop-Fan propulsion concept have been investigated in numerous propulslon and aircraft systems studies conducted by the airframe and engine manufacturers under NASA sponsorship (References 3 thru 9). These studies have shown that the inherent efficiency advantage that turboprop propulsion systems have achieved at lower cruise speeds may now be exte,ded to the higher speeds of todays turbofan and turbojet powered alrc_aft. By applying swept wlng/reduced dlameter technology to the design of prol=eller blades, and by achieving higher disc 1oadlngs through the use of a greater number of blades, It was found that the inherent fuel efflclencyof t_e propeller can be extended to speeds up to .85 Mn. The efficiency of the l_rop-Fan should a11ow aircraft to be designed that are 15% to 25% more eff$cIent than today's most technologlcally advanced turbofan powered airlines (l_mference I0).

Since 1975, Hamllton Standard has been deeply Involved v_lfl_, the NASA Lewis Research Center In the development of the Prop-Fan. I/_I_ re{ently, this effort utillzed a series of .622 meter (2 ft.) dlametermc_e_s which Incorporated differing numbers of blades as well as chamg_ _n blade shape.

These models underwent exhaust|ve testing in several wlm_ _mels at NASA and OF POOR Q'J_kb! _ '_ -L ORICINAL PAOE _.A,.,r_ A_'4D WHITE - PHOTOGRAPH FIGURE 3-1. I_OP-FAN CONCEPT 3.0 (Contlnued) United Technologies as wel] as on a NASA Jetstar acoustic research vehicle.

In these tests the targeted efficiencles were demonstrated, the source noise was characterized, and structural dynamics verlf|ed. Detailed descriptions of these tests and results have been the subject of numerous technical papers; a summary of which can be found in References II and 12.

Although the results of the aerodynamic performance and source nolse tests can be confidently scaled from model to product size, the structure of the solid homogeneous model blades Is so different from that envisioned for a product that extrapolation of model structural behavlor Is unacceptable. The verification of the structural integrity of a large-scale Prop-Fan then becomes the final major technical hurdle to be crossed before industry acceptance of the Prop-Fan as a viable aircraft propulsion scheme. This verification was initiated in 1983, when Hamilton Standard, under the sponsorship of NASA Lewis Research Center, embarked on a program to build and test a 2.74 meter (9 foot) diameter, 8-bladed Prop-Fan. Thls Prop-Fan was designated the SR-7L or the Large Scale Advanced Prop-Fan (LAP). The major elements of the LAP program are depicted in the summary schedule shown in Figure 3-2. Detail design and fabrication of the Prop-Fan components (blades, hub and blade retention, spinner, pitch change mechanism, pitch control and instrumentation system) was initiated early In 1983 building on a preliminary design conducted under an earlier contract. Various bench tests of each component then followed to verify key design characteristics.

Design, fabrication and test of an aeroelastlcally scaled .622 meter (2 foot) Prop-Fan model was included in the LAP program to obtaln an early assessment of the Prop-Fan's aeroelastic characteristics. This model has been designated SR-7A. Other objectives of the SR-7A testing Included the measurement of aerodynamlc performance and noise.

Testing of the SR-7L rotor under the LAP program included in-house whirl, static and high speed wind tunnel tests. Whirl Rig Testing was successfully conducted on the G-5 rlg at Hamilton Standard. The objectives of the test were to measure the stiffness of the blade retentlon system, evaluate the control dynamic characteristics of the blade pitch change system and determine the wear rate on blade actuation and retention hardware. Whirl Rig Testlng was conducted using stub weights to simulate the Prop-Fan blades. The stubs provided appropriate centrifugal loading but generated essentially no thrust or drag. The Static Rotor Testing, with SR-7L blades installed, was success- fully conducted at the Wright Aeronautical Laboratory at Wrlght Patterson Air Force Base in Dayton, Ohio (Reference 13). The goals of the Static Rotor Test were to measure the static aerodynamic performance of the LAP, assess stall flutter characteristics and investigate the structural behavior and integrity of the SR-TL blades for static operating conditions. The final component of the LAP test program, High Speed Wind Tunnel Testing, is the subject of this report.

{/1 I- Z LI.I 1.1.1 J r: 1.1.1 < E r _L rL < .-1 Z < ,-,I 1.,r, I,-.

o rl,.

W U

"ll

Z < < t_ < uJ < U _J < a. -i ,,.J I-,- _'1' w W c C U.@ ki D.

E _

O0 0 _ o

o.I=

ORIGINAL PACE BLACK AND WHITE FHOIOGRAPN" 3.0 (Continued) The LAP program is comPlemented by another NASA sponsored program, the Prop-Fan Test Assessment (PTA) program, which takes the large scale Prop-Fan (developed under the LAP program) and mates it with an Allison Gas Turbine supplied gas generator and gearbox to form a Prop-Fan propulsion system. The ma_or elements of the PTA program are depicted In Figure 3-3. Following completion of the Static Englne Test at Rohr's Brown Field Facl]ity, the quick engine change (QEC) nacelle will be mated to the wing of Gulfstream II aircraft which will ultimately serve as the f]|ght test vehicle for the Prop-Fan.

This report addresses the procedures and results of the High Speed Hind Tunnel Test conducted at the ONERA SI-MA Atmospheric _ind Tunnel in Modane, France.

This facility was selected for three reasons. First, it is capable of reaching high cruise Mach numbers. Second, it is sufficiently large (8 meter or 26.25 ft. diameter test section) to avoid excessive wall interference effects; and third, it has an existing model drive system. Although the power capability of the drive Is only about one fourth of what the Prop-Fan is designed to absorb, proper blade loading can be reached by running with a partial set of blades (eight, four, and two blade configurations). The specific objectives of the test are listed below.

To conduct a careful and controlled search for any evidence of classical flutter. Because of the greater air density of the :nd tunnel, it is possible to more closely approach the flutter threshold than at the 10,668 meter (35,000 ft) flight altitude. At design Mach number the wind tunnel operates at an effective altitude of about 4,267 meters (14,000 ft). Analytic predictions and tests of the SR-7A model strongly suggested that classical flutter would not be encountered.

To measure steady and unsteady surface pressure on the blade as well as overall Prop-Fan performance. One blade was Instrumented with 465 static taps (20 chordal at 13 radial stations on the camber side and |6 chordal at 13 radial stations on the face side of the blade) to obtain a complete pressure map. Another blade had 26 dynamic pressure sensors (7 chordal at the 35 inch station and 6 chordal at the 49 inch station, for each side of the blade) to assess unsteady effects. These measurements will provide bench_rk data for understanding the physlcs of transonic flow over the blades and for verlficatlon of analytic computer codes.

To determine the structural and aerodynamic response of the Prop-Fan to angular Inflow. Analysis of data from this simple, known angular inflow condition will slgnlfIcantly contribute to the understanding of Prop-Fan behavior in the more complex, airplane installed flow field.

LARGE - SCALE PROPFAN FROM LAP UNIQUE TILT NACELLE MOOIirlED TURBOSHAFT II b / \ II ENGINE GEARBOX /_ " _SYSTEM STATIC TEST AT ROHR BROWN FIELD

AIRCRAFT .OOEL TESTS \ / /

k; (.: ._.'_ " /'_ _._ CAB,,,ENVIRO MENT

1 T- 1.2f,_;_lN tillS" _ "-;-'_',_ EN-ROUTE NOISE L I _'I._JL_ - FLIGHT TESTING _ "_-a'mlJ_a_-,'_Imb _ INSTALLED OPERATING CHARACTERISTICS FIGURE 3-3. PROP - FAN TEST ASSESSMENT (PTA} PROGRAM ELEMENTS 4.0 PROP-FAN DESCRIPTION 4.1 General Description The Large Scale Advanced Prop-Fan shown _n Figure 4-1, is a 2.74 meter (9 ft.) diameter 8 bladed tractor type Prop-Fan rated for 4474 KN (6,000 HP) at 1698 RPM. To achieve the program objective of verifying large scale Prop-Fan structural integrity, a number of design requirements and goals were established as summarized in Figure 4-2. The requirements Include characteristics Judged essentlal to meeting the program objective as well as design features established from prior work. The goals, on the other hand, represent design targets and were judged less Important to the program objective. The LAP is designed to be mounted on a standard 60A spIined propeller shaft for an existing turboprop gearbox. It has a hydrau11cally actuated blade pitch change system and a hydromechanlcal pitch control that allows the Prop-Fan to operate In a speed governing mode or, with minor modifications, in a Beta Control mode, as was the case in this test. Beta Control mode operation was chosen to provide the operator the capability to select desired blade pitch angles while running. A pitchlock feature Is also incorporated in the actuator. This feature malntalns the propeller blade angle In the event of a loss of system operating oil pressure. The design of the actuator and control is based on proven technology used in Hamilton Standard's military and commercial propellers. A brief description of each of the major elements of the LAP as depicted in Figure 4-I, is presented below.

4.2 SR-7L Blade Features of the structural configuration of the SR-7L blade are shown in Figure 4-3. These Include a central aluminum spar which forms the structural "backbone" of the blade, a multi-layered glass-cloth-reinforced shell overhanging the leading and trailing edge of the spar, a nickel sheath which covers the leading edge of the outer two-thirds of the blade, and a non-operational integral heater In the Inboard leadlng edge area. Though the scope of the LAP testing never Included utillzation of the blade heaters, It was decided to Install the heaters to evaluate the structural response of a blade closely resembling that of a typical blade configuration. The remaining internal cavitles are filled wlth Iow-denslty rigid foam. The outboard portion of the spar is Intentionally moved toward the blade leading edge to increase stabillty by reducing overhung mass in the tip trailing edge, while at the same time increasing the Integrity of the leadlng edge from the stand- point of resistance to foreign object damage.

The blade design makes use of a NACA Series 16 airfoil outboard and a Serles 65 clrcular arc airfoil Inboard. Each blade has an activity factor of 227.3 wlth 45" of blade sweep at the tip. The blades were designed with predeflectlon so that the blades will assume the desired aerodynamic shape at the cruise operating condition (Reference 14).

=: BLADE CONTROL BLADE RETENTION ACTUATOR SPI LARGE-SCALE ADVANCED PROP-FAN FIGURE 4-I.

O_!GINAV p_v _Lf-,CK At, iD WHi-rE PHOIOGRAPH TECHNOLOGIES _ UNITED

DESIGN REQUIREMENTS

Configuration Diameter 2.74 meter (9 ft) Number of blades 6 • Design point 0.8 Cruise Mach number Altitude 10,668 meters (35,000 ft) 244 meterslsec (800 ft/sec) Tip speed Power loading (SHPID2) • Structural Integrity Flutter free over normal flight envelope (M <0.8) Stresses within allowable limits Overspeed tolerance Critical speed margins Safety features Leading edge projection Lighting protection Icing protection (installed but not operational) Overspeod protection • Reverse thrust capability

DESIGN GOALS

-78.6% AT M = 0.8, 10,668 meters (35,000 ftXCruise) • Net efficiency (isolated nacelle) -52.0% AT M = 0.2, SL (TO) • Noise .144 db overall sound pressure level Near field (design point cruise, max.

free field, 0.81)) •FAR 36 minus 10 db Far field • Stall flutter -None st 100% TO power and rpm; M = 0- 0. 2 • High speed (classical) flutter -None over extended flight envelope, (M __< 0.85) 105% max operating speed .120% max operating speed - no yield • Over speed limit (hub, blades, -141% max operating speed - no failure blade retention) Foreign object damage -No damage to pdmary blade structure Minor-Birds up to 4 oz -Some loss of matedal or airfoil distortion; Moderate -2" Hail; Birds to 2 Ib operate at 76% power for 5 minutes -Some loss of matedal or airfoil distortion; Major-Birds up to 4 Ib maintain ability to feather • Blade life .35,000 hr - replacement with scheduled malnt.

-50,000 hr - meantime between unscheduled removal FIGURE 4-2. DESIGN REQUIREMENTS AND GOALS SUMMARY z FIGURE 4-3. FEATURES OF THE SR-7L BLADE CONSTRUCTION 4.2 (Contlnued) Although some improvements in sweep/stress/stability trade-offs were predicted through the use of advanced composites, it was decided not to include these In the final blade design. Their use would requlre the development of new manufacturing technology, both in terms of suitable construction methods and processes, and lengthy development of design allowables to reflect the manufacturing process.

It was felt that the scope of the program would be best served by utilizlng the service-proven combination of an aluminum spar enveloped with a fiberglass shell for which processes and stress allowables are well known.

4.3 Pitch Change Actuator The pitch change system Is comprised of two components, a pitch change actuator and a control. The pitch change actuator Is the prime mover for blade angle change and Is located within the Prop-Fan hub as shown in Figure 4-4. Its primary components are an internal stationary piston, a translating outer cylinder with an integral yoke to engage each of the blade trunnlons and a pltchlock and servo assembly which contains a four-way metering beta valve assembly, a pitchlock screw, a ground adjustable low pitch stop, a servo piston and a ball screw to drive the pitchlock screw and beta valve (Reference 15).

As stated earlier, the LAP was modified to operate In the Beta Control mode throughout all phases of the wind tunnel test. This was accomplished through the use of an electromechanlcal controller (D.C. motor and gearhead) mounted on the front of the dome as shown in Figure 4-5. The controller provides the rotary input to the pitchlock servo directly, replacing the rotary input from the half area servo. In order to give the D.C. motor full control of the blade angle, the servo and ballscrew must be disconnected from the pitchlock screw by removing the quill shaft. This disables the control signal and allows the motor to work directly on the pitchlock screw without having to fight the servo output. This permits the operator to remotely position blade angle as deslred while running.

The actuator was designed to present state-of-the-art technology and low development risk technique that has been used on a number of existing propeller systems. The design uses mostly steel for the load carrying members and all surfaces subject to sliding seal wear are chrome plated to increase durability. The actuator was designed to conservative stress and deflection levels to minimize development effort while maintaining a reasonable but not minimum weight.

The pltch change mechanism was designed such that all malfunctions will either cause the system to pltchIock or go to Feather. An additional safety Feature on the LAP is a ground adjustable low pitch stop. This 11mlts the minimum blade angle under all circumstances.

LAP PITCH CHANGE ACTUATOR FIGURE 4-4 ORIGINAL PAGE AND WHITE pI_OTOGRAPIq

BLACK

± _.= u o a-(n m Z < -r U ILl :E ..J n, Z (J u !- ;[ IJJ > m Q I- o W 4.4 Pitch Control To minimize cost and development time, It was decided to utilize a modified version of an existing turboprop propeller control. Based on the type of engine and gearbox planned for use with the LAP, the control selected was a modified 54460 control (Figure 4-6). The 54460 hydromechanlcal integral oii control, which Is currently In use on the Grumman E-2 and C-2 aircraft, was modeled after the Lockheed C-130 and P-3 controls. Since the first production unit in 1956, there have been over 12,000 built and they have logged over 80,000,000 operating hours.

The primary func_ion of the LAP control, as modified for wind tunnel testing, was to generate the hydraulic pressure necessary to assure proper actuator operation. Two gear type pumps located in the stationary control and driven by the Prop-Fan shaft provided the system hydraulic pressure.

4.5 Hub and Blade Retention The LAP hub assembly forms a seml-rlgid llnk between the blades, which provide the thrust, and the engine shaft, which provides the torque (Reference 16). The hub and tailshaft is a one piece forged component which is carburlzed, heat treated and machined (Figure 4-7). A single row ball bearing retains each of the eight blades In the hub, while the tailshaft secures the Prop-Fan to the engine shaft through two cone seats that are preloaded against each other by the Prop-Fan retaining nut. The hub also forms the support for the pitch change actuator system, the control and the spinner.

The retention transmits the loads from the blades to the hub while accommodating changes in blade pitch. The single row ball bearing retention provides ease of maintenance by allowing individual blade replacement without disassembly of the hub. It has a through hardened Inner race which seats against the aluminum blade shank and an outer race which Is integral with the barrel. The outer race Is carburlzed to achieve the hardness necessary to support the ball loads. The balls are kept apart from each other by an elastomerIc separator. The rotational speed of the Prop-Fan keeps the retention submerged In oll which is contained in the hub by eight blade seals (Reference 17).

4.6 Spinner The LAP spinner and rear bulkhead assembly Is essentially a reinforced flberglass/epoxy shell, supported by the hub and actuator, and Incorporating an aerodynamic shape to facilitate proper alr flow around the blade roots (Figure 4-8). Its primary function is to insure proper Prop-Fan aerodynamic performance. The rear bulkhead, which mounts on the rear of the hub arms, Is the main structural support for the spinner and provides a mounting surface for much of the instrumentation hardware in the rotating field.

!

!

n -r _I D_ I- Z U D.

..I lg I_CK Abed WHITE Pi-_OTOGRAP_},:I FIGURE 4-7. SR-7L HUB OP.. IGIF-'#.L P/_.OE' BLACK A _'_n, -,,., '.,f.,'r4t"t-E -.I ,,- -._.--,D,, Du,_ ,...,,,j,...,,;.., FIGURE 4-8 SR-7L SPINNER 21/22 r_ 5.0 INSTRUMENTATION DESCRIPTION 5.1 General Description In order to accomplish the objectives set forth in the Plan of Test, three separate Instrumentatlon arrangements were employed to gather and record the desired data. For the structural dynamic and aerodynamic performance tests, the system described in sections 5.1.I and 5.1.2 was used. For the blade surface steady pressure test, a system description is provided In section 5.2 and for the unsteady pressure test, the system used is presented in section 5.3. Figure 5-I provides a summary listing of all the instrumentation, rotating and stationary, and indicates whether It was provided by H.S. or ONERA. Additional detailed descriptions of the key data gathering devlced required for specific phases of the test are presented in the appropriate Test Procedure section of this report.

Common to all three instrumentation arrangements was the frequency modulated multlplex electronic data acquisition system (DAS), as shown in Figure 5-2, and several Prop-Fan system diagnostic monitoring devices.

5.1.1 Electronic Data Acquisition Systems (DAS) The electronlc data acquisition system for the LAP provided the capacity to transmit 33 channels of information from the electronic measurement devices on the rotating slde of the Prop-Fan to the data collection and monitoring equipment in the stationary field. This was accomplished by employing an eight rlng platter-type sllp ring assembly which provided the electrlcal Interface for the DAS between the rotating Prop-Fan assembly and the non-rotatlng control.

As illustrated in Figure 5-2, three rings were utilized to transmit data to the stationary field. One of the three carried the output of a potentiometer which was mounted as shown in Figure 5-3 and used for monitoring blade angle position. The other two rings transm|tted the remaining 32 channels of information which consisted of signals from two miniature pressure transducers for monitoring actuator high and low pitch pressures and combinations of up to 30 blade strain or unsteady pressure signals, as determined by the particular test being conducted. Transmittal of these 32 signals on only two sllp rings necessitated the use of FM multiplexing. The DC signals from the strain gages and pressure transducers in the rotating field were divided into two groups of sixteen. The signals were then converted to frequency modulated signals by two groups of voltage controlled oscillators. Each group was then multlplexed by a mixer, allowing thirty two channels to be transmitted through two sllp rings. The two groups of sixteen channels were each detranslated in the stationary field to four groups of four multiplexed channels (IRIG Standard IA thru 4A). Each set of four channels was recorded on one track of a standard Honeywell IOl tape recorder. Simultaneously, eight discriminators were used to demodulate any two groups of four channels for real tlme monitoring of data. One discriminator was tuned to the center frequency of each channel.

_G PAGE H.S. INSTRUMENTATION ROTATING: HIGH AND LOW PITCH OIL PRESSURE MEASUREMENT BLADE ANGLE MEASUREMENT BLADE VIBRATORY STRAIN MEASUREMENTS BLADE SURFACE STEADY AND UNSTEADY PRESSURE MEASUREMENTS STATIONARY: CONTROL SUPPLY PRESSURE MEASUREMENT CONTROL SUMP OIL TEMPERATURE MEASUREMENT HEAT EXCHANGER _P MEASUREMENT CONTROL AND SCANIVALVE ACCELEROMETER MEASUREMENTS INSTRUMENTATION PROVIDED BY ONERA (ALL STATIONARY) TEST RIG INSTRUMENTATION: RIG ACCELEROMETER MEASUREMENTS PROP-FAN RPM PICK-UP RIG BEARING TEMPERATURE MEASUREMENTS BALANCE AND TORQUEMETER MEASUREMENTS (FIGURE 6-I0) CENTERBODY PRESSURE MEASUREMENTS (FIGURES 7-I0, 7-11) SPINNER BULKHEAD PRESSURE MEASUREMENTS (FIGURE 7-]2) TUNNEL INSTRUMENTATION: STATIC PRESSURE MEASUREMENTS: - 4 METERS UPSTREAM OF THE PROP-FAN PLANE OF ROTATION - ADJACENT TO THE PROP-FAN PLANE OF ROTATION STAGNATION PRESSURE MEASUREMENTS AIR DENSITY MEASUREMENT STAGNATION TEMPERATURE MEASUREMENT OF FLOW STATIC TEMPERATURE MEASUREMENT OF FLOW FIGURE 5-]. H.S. AND ONERA INSTRUMENTATION LIST u I b- < !

_Jl _J

!

W l- >- z b- < I- z b.I n, 1- (/) z < -i W

i-l-l-l-l-l-l-i-

FIGURE5-3. LAPBLADEPITCHANGLEMEASUREMENT SYSTEM

5.1.1 (Continued) The FM electronic instrumentation system provided inherent noise immunity for data transmission. The frequency response for the system was DC to lO00 HZ.

Overall accuracy of the system was ±3% RSS. Time correlation between channels was +13.8 mlcroseconds.

The electronic Instrumentation system allowed for up to ten blade shell or shank gages to be installed on any blade, though a maximum of 32 gages could be active at any one time. Straln gaged blades were closely monitored during flutter and critical speed testing as well as measuring strains at various operating conditions. The blade shank strain gages were employed to measure blade bending moments.

A total of sixteen gages could be selected from blades one through four and an additional sixteen gages could be chosen from blades flve through eight.

Selection of the desired combination of gages was accomplished using eight programmable connectors mounted on the Prop-Fan hub. Programming of the connectors required using Jumper wires to connect the sockets of patch boards in the connectors. The bridge completion circuits for the strain gages were located on circuit boards in the blade cuff.

Monitoring of instrumentation during the test was accomplished with an oscilloscope, a spectrum analyzer and a vlsicorder. The oscilloscope permitted a time domain display of four channels simultaneously. The spectrum analyzer provided the capability to dlsplay any one channel in the time or frequency domain. The analyzer also had transient capture and playback features. The vlslcorder provided a hard copy plot of instrumentation signals versus time.

5.1.2 Prop-Fan Diagnostic Monitoring Instrumentation In addition to monltorlng and collecting data from the rotating fleld, a number of parameters intended to provlde protection for the Prop-Fan system were measured and digitally dlspIayed by the stationary field instrumentation.

These included control sump oil temperature, control supply oil pressure, differential pressure across the heat exchanger and control/Scanlvalve accelerometer measurements. The control sump temperature was measured by a thermocouple installed inside the oll draln port of the control. Control supply oil pressure was measured by a transducer Inslde the control. This transducer slgnal was transmitted via an existing connector on the control.

The heat exchanger differential pressure was measured by a AP transducer connected across the o11 exlt and return ports on the aft face of the control.

The control vibration was measured by two accelerometers mounted on the propeller control houslng. One accelerometer was oriented to sense motlon in the horlzontal dlrectlon and the other to sense motlon in the vertlcal 5.1.2 (Continued) directlon. Additionally, during conduct of the steady pressure test, an accelerometer sensing vertical motion, was installed on the Scanlvalve fairing assembly. The "once per revolution" signal was orlglnally planned to be provided by a pickup mounted on the control and triggered by a target on the rotating propeller bulkhead. However, the system eventually used was provided by an eddy current proximity probe targeted on a 59 tooth wheel (on a 60 t(_oth basis) mounted on the propeller shaft. The rotational speed was averaged over ten revolutions.

5.2 Steady Pressure Measurement System In addition to employing the instrumentation systems described In section 5.I.I and 5.1.2, a specially designed pneumatic instrumentation system was used to collect and measure blade airfoil surface steady pressures. Thls system consisted of a speclally fabricated blade with rows of pressure taps Installed at thirteen radial stations and a scanlvalve mounted on the nose of the Prop-Fan. (See Figure 5-4.) The pressure taps were connected to the scanlvalve by 36 capillary tubes run along the acutator dome.

The scanivalve Instrumentation system provided 36 channels for transmitting steady pressure data from the surface of the blade to the stationary field..

The scanIvalve itself consists of a rotating and a stationary portion. The rotatlng portion was attached to the front of the actuator dome. The radial tubes from the steady pressure measurement blade were connected to the rotatlng portion of the scanlvalve. Each tube was connected to one channel of the scanlvalve. The statlonary portion of the scanlvalve contained a pressure transducer which monitored one channel of the scanIvalve at a time.

The stationary portion of the valve protruded through the leading edge of the propeller spinner and was restrained against rotation. Switching of the scanlvalve channels to be monitored by the transducer was controlled by a pneumatic signal requiring a clean alr source of 150 psi. The scanning rate was adjustable from 0.l to lO seconds per channel. The scanivaIve was enclosed In an aerodynamically shaped fiberglass fairing to maintain a well behaved inflow to the Prop-Fan. The umbilical, which connected the scanlvalve wlth the control and monitoring equipment outside the tunnel, was enclosed In a conduit with an airfoil shaped cross section. This also minimized disturbance of the inflow to the Prop-Fan.

On 11ne monitoring of the scanlvalve pressure transducer sIgnals was accompllshed uslng the digital readout from the scanlvalve controller. In addltlon to recording the measured pressures on magnetic tape, on-site data manlpulatlon and prlnt-out generation was provlded by a Fluke DA computer coupled wlth a plotter and dot matrix printer.

OF., I_0 OR QdAL_I ,Y n \ O_ (/1 IX 5.3 Unsteady Pressure Measurement System Collection of blade surface unsteady pressure data was accomplished by utilizlng a specially instrumented blade, coupled with the instrumentation arrangement described earlier in sections 5.l.l and 5,1.2. Twenty-six high frequency response pressure transducers were |nstalled in two rows on each side of the blade. The unsteady pressure signals were transmitted from the rotating to the stationary field through the FM multiplex electronic data acquisition system. The signals were monitored on the four-channe] oscilloscope and recorded on the IRIG tape recorder.

L 6.0 TEST FACILITY DESCRIPTION 6.1 Wind Tunnel Description The LAP High Speed Rotor Test was performed in Chariot No. 3 In the SI-MA Large Wind Tunnel at the Modane-Avrleux Aerothermodynamlc Test Center operated by the Office National D'Etudes et des Recherches Aerospatlales (ONERA), France (Figures 6-I and 6-2). The tunnel is a continuous, closed loop, single return, atmospheric facility incorporating bleed slots In the tunnel walls to a11ow air exchange capability and tunnel circuit screens to mlnlmlze turbulence, as shown In Figures 6-3 and 6-4. Tunnel cooling is accomplished by alr exchange wlth outside air of up to 20% of the test section mass flow rate. The tunnel Is driven by two coaxial counter-rotating Pelton water turbines from a water supply wlth an 865 meter head, sufficient to generate a maximum of 88MW (I00,000 HP). The test section velocity can be continuously varied from 8 m/sec to Mach 1.0, well encompassing the range for this test. The curves of Figure 6-5 glve a representative sample of the tunnel driving power and Reynolds number as a function of test section Mach number. Tunnel pressure altitude varies from approximately 1,100 meters (3,600 ft.) at low speed, up to 6,000 meters (20,000 ft.) at Mach l.O. At Mach 0.8 the tunnel operates at an effective pressure altitude of approximately 4,260 meters (14,000 ft.). Stagnation temperature ranges from -20°C to +60°C, depending on the external ambient temperature and the Mach number in the test sectlon.

The tunnel has three Interchangeable test sections or chariots which are positioned in the aerodynamic clrcult. Each of the chariots contain a different test section shape and size and are speclallzed for a wide variety of test capabilities. The use of Individual chariots a11ows a test to be conducted in the tunnel while the next test Is being set-up In another chariot located In l of 2 adjacent mounting stations. The LAP was installed In Chariot No. 3 which has a propeller test rlg permanently mounted In its test section (Figure 6-6). The test section Is 14 meters long, has a circular cross section with a diameter of about 8 meters and non-permeable walls.

Several lateral fillers or Inserts were added to the test section specifically for this test to assure the desired range of Mach numbers could be attained.

They serve the addltional purpose of area compensating for the propeller test rlg blockage generated by the 970 mm diameter center body. A more detailed description of the tunnel facilities is contained in References 18 and 19.

6.2 Test Rig Description The Prop-Fan was Installed on the test rig as shown In Flgure 6-6. The power source for the Prop-Fan drive system Is twln Turbomeca gas turbines driving a common gearbox, enclosed In a streamlined pod located approximately ? meters downstream of the Prop-Fan plane, as shown In Figure 6-7. The engines were rated for a combined maximum power of 1000 KW (1341) HP at standard conditions (59°F, 29.92 In. Hg.) as depicted In Figure 6-8. Velocity measurements made upstream of the engine Inlets Indicated no disturbance at the Prop-Fan plane due to the operation of the engines. The gearbox was a \ o 50 lOOM FIGURE 6-1. MODANE - AVRIEUX AEROTHERMODYNAMIC TEST CENTER ORtGfNAL PAGE BLACK AND WHITE PHOTOC_RAPH FIGURE 6-2 S1-MA LARGE ATMOSPHERIC WIND TUNNEL _ 15_.M "-I

/

I _. T.oT,_IR L,- - iil_l

,0M E×..osT I ! WORKING CO -AXI'L_I

I I -_" .... '-I _Ec_ION _--N-il !, _,,._',,

! i//I I_ _-_-J_,.._t_-I_ , w , il_ _r

---t-_.-l-.-I_-l-l-.:l_tfl-.-I- , -_i--._ IIl-'b,_J

I// I I IL I1.,"1 'i'T'_-FI i , L I". _! !','_',!

. .. ,_" ! ,. IT !r l lltIH i I!l]-._.:_l

I s ,,,ov.o,.E =,-,.R,OTS _ I _!i |i WATER "_-_.-_ -- I AIR INLET I

F__l_-I i!_

L.__ '-° ..... _ :li WATEr' PIPE . FOR ENGINES !I; AH = 865M FIGURE 6-3. S1-MA WIND TUNNEL SCHEMATIC _4 p ,_ _, ,_,

ORIGINAL

;-; _ ._

BLACK

AN O WH ITE

PI-.iOTOGRAPH FIGURE 6-4. S1-MA WIND TUNNEL NOZZLE INLET "I1B_SE'¢TION 3.5 e (MWJ lOO 8o 6 60 4o 2o MACH 1 0.5 NOTE: THE CHART ABOVE IS FOR A REFERENCE CHORD LENGTH OF 0.7 M TUNNEL DRIVING POWER AND REYNOLDS NUMBER AS A FIGURE 6-5.

FUNCTION OF MACH NUMBER Z t-, r_ < -I- i.i z

o)

F- ILl -I-- I-- Z i Z I

f

l- < -I -i < I- Z Z < O_ w bJ tg m" C_ z z z <_ Iil I- !

z 0R!0!NAL PA__ BLACK AND WHITE PHOTOGRAPH = .

- O_--POOR QUALITY FIGURE 6-7. PROP-FAN DRIVE SYSTEM ENGINIES_I_II_IF_.ARBOX

Kwi

OPERATION LIMITED TO 5 MINUTES 1oo0 ..

(MAX, CONTINUOUS) POWER ._'_ ___'_"_'_ _NORMAL OPERATING REGION !:._j ....',_;..._,_._'i b ',-_'7;._:_ ";'_', 7 :,._,,. -_,-"'_': ",_";: _.:.' _':'_ _.;_;_ t_-_ • _,; "_'Z,; ":. :" :".'4_'; ........... ,,_....._._..;1";$>,,. "_ ............ ,.7, .- .'-,_; ," ,',-'_ ,.. ":;'_" "'_'"""%% _;:'_!".,_'_ ".::_!,.,::%_._:'_._._,_._'Z'._<_s; "z. ....._.._. ..... . ..., ._. ,..

0 500 I000 1500 2000 22.00 2460 PRPM MINIMUM PERMISSIBLE PROP-FAN SPEED FIGURE 6-8. PROP-FAN TEST RIG POWER LIMITATIONS 6.2 (Continued) speed reducer providing a reduction in RPM of approximately 3:1 from the engine to the Prop-Fan. Power was transmitted to the Prop-Fan from the gearbox through a long drive shaft and a balance (Figure 6-9). The drive shaft housing was supported by rods and damping cables attached to the tunnel walls. The particular rod and cable arrangement used In each test was dependent upon the type of test being run and the number of blades installed. These arrangements were established during a pre-test vibration survey conducted In the tunnel by ONERA. Universal joints at each end of the drive shaft allowed the Prop-Fan to be operated at various inflow angles relative to the flow through the test section.

A balance and torquemeter, Installed in the drive system, allowed the thrust, torque, sides forces and bending moments acting on the Prop-Fan to be measured. A diagram of the balance is shown Is Figure 6-10. Forces and moments are transmitted through the balance by six strain gaged elements whlch are referred to as dynamometers. Thrust, yawing moment and pitching moment are transmitted and measured by the three axial dynamometers (TI, T2 and T3). Balance frlctlon torque and the lateral and vertical forces are transmitted and measured by the three tangential dynamometers (Zl, Z2 and Y).

The torque supplied to the Prop-Fan is computed by subtracting the measured friction losses in the balance from the torque measured and transmitted by the torquemeter. A11 six dynamometers have a capacity of 2000 daN. Two flexible couplings decouple the transmission of the torque from the thrust measurement with one of the couplings being fitted with a gage bridge (torquemeter).

A stationary aerodynamic fairing or centerbody was installed around the drive system. This was also commonly referred to as the minimum body. The minimum body provided a downstream extension of the aerodynamic contour of the Prop-Fan spinner and was designed to reduce the air veloclty passing through the root portion of the Prop-Fan rotor. At the inner blade radii the combination of thick airfoil sections and the large number of blades presents significant blockage to the flow, which could result in choking if the velocity was not moderated. The minimum body Is approximatly 35% of the diameter of the Prop-Fan rotor. A description of the minimum body pressure tap arrangement is presented In Section 7.0.

6.3 Test Riq Modifications The LAP Prop-Fan High Speed Wind Tunnel Test was completed following the incorporation of several structural improvements to the propeller test rig drive system. These modifications, as shown In the comparison of Figures 6-II and 6-12, were added to preclude recurrence of a cone seat fretting corrosion problem encountered early In the testing, which led to an unanticipated interruption of the test. In addition, the modifications in genera! provided a more structurally sound rig design, better able to handle the loads Imparted by the Prop-Fan.

w !-- > w > n, r_ Z < I, n, ID w It.

U l- r IJJ "l- u u z < -I In Z < I.- 1.1.1 > k9 I I I I- Z < _k "7 _D IJ.l l.l..

i

Z < m U.

Z I- w I,- Z m ,2 "r ID W w z m Z < \ OU U.

z _z < b_.j \ U >< \\ (2b.

n, c: 0 1- b_w Ou I- ILl e,i _D W < )" k- b. Z < bJ -r _.

u) Id _. r_ 0 l" r_ _j U bI bl Z L 0 6.3 (Continued) The most significant modification to the drive system was the elimination of the cone seat retention configuration on the aft end of the prop shaft. It was this area which experienced the severe fretting corrosion problem early |n the test. The new design provides for a one piece prop shaft mounted in a substantially enlarged forward test rig housing incorporating increased capacity cylindrical roller bearings. Since a safe structural operating envelope for the Prop-Fan had been defined earlier in the testing, the balance was removed to provide room for the test rig structural improvements.

45/46 r 7.0 SPINNER AND CENTERBODY DRAG MEASUREMENT 7.1 Test Rig Vibration Survey Before testing was conducted In the SI-MA wind tunnel, the whirl flutter stabillty of the Prop-Fan/test rlg system was evaluated. Stability of the Prop-Fan/test rlg was a major concern because the LAP was the largest assembly ever Installed on the test rlg.

The Prop-Fan was assembled and installed on the propeller drive system in the tunnel test section in preparation for the vibration survey. To e11minate the possibility of damaging any of the actual LAP blades during this survey, blade stubs were installed in place of the blades. Additlonal weights were hung from selected blade stubs to account for the weight difference between the stubs and blades. The stubs were shimmed to minimize any movement in the blade retention during the shake test.

The survey was conducted using two electromechanlcal shakers positioned at various locations along the axis of the mlnlmum body. Data was obtained using two accelerometers mounted on the test rlg force balance Just aft of the Prop-Fan. The data was used to determine the elgenfrequencles of the various rlg vibration modes in the horizontal and vertical planes.

The results of the survey confirmed the need for an unsymmetrical rlg support structure which pre-test analytical whirl flutter stability studies had suggested. Though the vibration test demonstrated that the 4 bladed SR-7L would operate whirl flutter free, it was agreed that rig damping characteristics would be closely monitored during all test envelope expansion.

In order to incorporate the maximum test rlg resistance to whirl flutter onset, several different rlg support structure conflguratlons were employed, based on the number of blades Installed, as depicted In Figures 7-8, 8-I, 8-2, lO-I and ll-l.

7.2 Wlnd Tunnel Corrections Immediately following completion of the test rig vibration survey, ONERA conducted the wlnd tunnel calibration test. The purpose of this calibration was to remove the effects of the presence of the test rig on the measured Mach number. The ONERA approach was as follows. First, a compressible flow calculation was performed to establish the axial velocity distribution at the working radius of the blade (.6R) through the plane of rotation of the Prop-Fan. Next, two independent axial velocity surveys were made at the same radial location, utllizlng a speclal pressure tapped plpe and scanlvalve system. The results of these surveys are shown in Figures 7-) thru 7-6. The uncorrected Mach number data represents actual measurements _alken, while the corrected Mach number data is the difference between the ¢a_¢ulatlons and the measurements of Flgures 7-I thru 7-6 added to the wall leach m_:mber measured 2 meters (6.56 ft.) upstream of the Prop-Fan plane of rotat_om. Thus, the corrected Mach number data represents Moo= .354 ;RRECTED MACH NUMBER r,, @ ¢ IO ¢o :E D Z 3: U < I X .I < UNCORRECTED MACH NUMBER ..,I c

I

500 MM 1000 MM |o00 MM 500 MM BODY LENGTH, MILLIMETERS FIGURE 7-1. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAM MACH NUMBER .354 n, n., MACH NUMBER en Z "!"

u J u .I CORRECTED MACH NUMBER I I 1000 MM !000 MM 500 MM 0 500 MM BODY LENGTH, MILLIMETERS FIGURE 7-2. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAM MACH NUMBER .615 a, tO v a, m _r Z U < :[ .J < MACH NUMBER ,1 .73 .72 MACFI NUMBER !

I .71 , !000 MM 500 MM 0 _ MM ! 000 MM I BODY LENGTI'L MILLIMETERS FIGURE 7-3. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAM MACH NUMBER .72I 5O Moo,, .754 I ,8O I .79 .78 .77 CORRECTED MACH NUMBER I .76 _, .75 .74 E io ""*__UUNCORRECTED MACH NUMBER ¢ .7_ Id :i z = u < t" .J t'*" u ,J I 1000 MM 500 MM 0 500 MM I ooo MM BODY LENGTH, MILLIMETERS FIGURE 7-4. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAM MACH NUMBER .754 .SS .84 .83 MN .788 .82 .81 U} m .80 :E Z 'I" .79 u < :E .J < .78 _J ..I ,77 .76

I

q.

, //, I 1000 MM 500 MM 0 500 MM 1000 MM I BODY LENGTH, MILLIMETERS FIGURE 7-5. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAM MACH NUMBER .788 ,gZ ,90 ' Moo" . I .89 829 .88 .87 r,* CORRECTED MACH NUMBER .85 m _E .86 I Z .84 Z o < X ,J < u

"" I -?---J

.J

Fo.oo-_o,_o..o,.o.m

,81 _ _"_'_ '" .80 1000 MM FIGURE 7-6. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAM MACH NUMBER .829 7.2 (Continued) the Influence of the tunnel walls on the axial velocity distribution. Note there Is very little difference at the lower Math numbers between the corrected Mach number data and the uncorrected Mach number data as measured at the 2 meter wall location. In addition to the velocity calibration described above, ONERA used the Prandlt-Young correction as discussed in Reference 20 to further adjust the velocity data before computing an advance ratio, 3. This correction, as defined below, is a compressible correction to the velocity to account for tunnel wall interference.

l _4 _1 II l w V 2(1 + 2 _4)I/2<1 - MN2) I/2 or Vcoe m 1 - Vmeas 2(I + 2 _4)I/2(I - MN_) I/2 where THRUST '1:4 p x DISC AREA x V 2 and PROPELLER DISC AREA TUNNEL CROSS-SECTION AREA Note that this correction Is largest where the thrust Is largest and the velocity the smallest, _.e. low Mach number testing.

7.3 Test Rig Corrections 7.3.1 Test Objectives 7.3.1.I To determine the aerodynamic drag on the Prop-Fan spinner as a function of Mach number for Mach numbers ranglng from .2 to .85.

7.3.1.2 To determine the aerodynamic pressure drag on the test rig centerbody as a Function of Mach number for Mach numbers ranging From .2 to .85.

7.3.2 Test Procedure In preparation for the splnner/body tare test, the test rlg and associated support structure were set up as depicted In Figures 7-7 and 7-8. In order to isolate and evaluate the affects of the presence of the spinner and center- body In the flow fleld, the LAP blades were replaced with blade stubs, whose external contours were machined to match that of the spinner (Figure 7-9).

As defined in Figures 7-I0 and 7-II, the Forward end of the centerbody serves as an extension of the external contour of the spinner, and was designed to alleviate compressibility losses in the blade root section by reducing the air velocity passing through the central portion of the Prop-Fan rotor. At the inner blade radii, the combination of thick airfoil sections and the large number of blades presents significant blockage which could lead to choked flow. The moderation of velocity caused by the centerbody reduces the possibility of choked flow occurlng at the high subsonic Mach numbers at which the LAP operates. Figures 7-I0 and 7-II also define the location of the centerbody surface static pressure taps. There are four rows of twenty-nlne taps per row positioned a]ong the length of the centerbody surface. The rows of taps were spaced 90 ° apart, clrcumferentially.

Figure 7-12 defines the location of an additlonal set of static pressure taps In the space between the spinner rear bulkhead and centerbody. These taps consisted of four rows, ten taps per row, extending radially outward, located at the centers of equal areas. Each row was oriented 90 ° apart from another.

Collection of the static pressure measurements from the locations described above, while varylng tunnel Mach number, was accomplished u_ing a scanivalve system. Splnner/centerbody drag data was collected at a total of 42 test points. All of the data was collected at zero degree Inflow angle.

Table 7-I lists the Mach numbers at which data was collected.

_g FIGURE 7-7. SPINNER/BODY TARE TEST .... v'_Pt_i _- PHO'i O&RAPI_I _tALANCE GEOMETRICAL CENTER BLADE PLAN _ AXIS WIND TUNN___ ._

vO -__

513.4"-_ ROD (_:35 ROD L:47Z0 _ _1 ._:26 ROD

SiL._-,o,o _ _::':,o

_o tTi_o B c

NOTE: ALL DIMENSIONS IN MILLIMETERS _) = DIAMETER L = LENGTH FIGURE 7-8. TEST RIG SUPPORT STRUCTURE FOR SPINNER/BODY TARE TEST 5?

< Id _-> _0 Z< r m _J v < m m_ Z < E_ b.

=_XXXX (1.

_ .zzzzz r,, h.

< w a _>_ PO_ w_ZZZ ;/ o_.

v N _N__ A Z Z w o I'- I- W L) U i.i Id U3 Q m 6 ........... ¢ < z < U w N Ii = ! 0 N <

IL

Z m

I1: 9

ta I- < Z LJ Id u O .J W Q.

<

It

!- 1.1.1 n,'

i t

(/'j ld N n,, Q.

/

W U < I.I.

n,'

t

IE IE >.

N C]

©C

O m n,, b.I I- E In Z b.l U bJ n,, (3 m k.

_9 m_mm_mmm_mv Z z m i-.

H- U U _J Ld (4 l- ]-.

b.

1.6 < < == _oo_oooo_o::::::: Q z m r_ r_ :E I- (/) Z u F_ I H- : m v" _" =', m _ _= _" N N N N N N N N N N F_ U o N -I N _L < hi T_ E F_ ,¢ Z U o o I- U < F_ -r< 1-.

U 0 gl U I,Ll i.-r < F_ < < @ b.

O F_ b.

m F_ 1:1 Z F_ U w I- u) F_ Z Z a Ld r_ b, X ). 0 F_ i m r_ w h z w u w utZ I-u <0 I-.l (_ NNNNNNNNNIq I- i: ..........

Z o I- < u .J Q.

< I- W ffl Ul I1.

U m I- < < U,I

0 0

_1 ILl Z Z E (I) m r_ _m I,d 0 N IZ

\

N m ___ m. _. _. m. m ?.

--me, iN N N _l (q (q eq _ m | o o 0 ooo0 00_ I (3 b.

Table 7-I. Splnner/Centerbody Drag Test Points SPINNER/CENTERBODY DRAG CORRECTED TEST MACH PT. NO.

25O 8OO .789 .739 .686 .638 •590 • 494 •447 2"/2 •348 .244 .201 .201 .201 .201 3OO .298 3O8 .298 .298 .299 31l .298 .201 .201 .201 .201 .201 7.3.3 Discussion and Results 7.3.3.1 Spinner Draq As depicted in Figure 7-13, the spinner drag (Ds) was measured directly from the axial force measured by the balance (Fs) with corrections for the back pressure force (FBp) and the losses due to thermal effects in the flexible coupling (FTH).

Ds = FB - FBp - FTH The axial force applied to the propeller shaft (F,) was measured by three dynamometers T,, Tz and T3 In the test rig balance as depicted In Figure 6-I0. The losses (FTH) due to the thermal effects of the flexible coupling were measured by strain gages bonded directly to the flex coupling.

The back pressure force (FBp) iS the result of the difference between the free stream static pressure (Po> and the integrated pressures fn the space between the rear of the spinner bulkhead and the face of the centerbody (PN), where aN IS an area weighting factor and A, Is the spinner base area.

4O FBp = ( _ (aNPN) -- Po) As N=I As - .507 m z Figure 7-14 presents the spinner drag coefficlent (Cso) as a function of corrected Mach number (McoR). The spinner drag coefficient Is given below and Is expressed In terms of the spinner drag force (Ds), dynamic pressure (qo) and projected spinner base area (As).

Ds CsD qoAs As = .519 m Since the spinner drag data was corrected for the back pressure effects, it represents the axlal components of the forces appl|ed by the static pressure acting normal to the spinner surface and the spinner boundary layer wall stress acting on the wetted area of the spinner. A quadratic equation was fitted to the spinner drag data, as shown below and on Figure 7-14, and was used to compute spinner drag corrections to measured thrust during Prop-Fan performance testing.

Cso - 2.7699M" - 5.3225M 3 + 3.7727M z - 1.1215M + .2398 co. co. co. co.

DCBT •",.'OE.E.U..--T_, / OS'_, _°P _T.

(FTH) DCBT " f(Ps -- PO) dA FLEXIBLE COUPLING LOSSES A. SPINNER/CENTERBODY TARE TEST (WITHOUT BLADES) TApP / / DCB BALANCE j MEASUREMENT (FB) B. PROP FAN TEST PT. (WITH BLADES) FIGURE 7-13, PROP-FAN TEST RIG FORCES 1,1,1 m Z L_ < r_ r_ U > Z u < r_ m _J o u 7.3.3.1 (Continued) The spinner drag coefficient, as expected, increased with Mach number for Mach numbers in the range from .3 to .85. However, an unexpected decrease in drag coefficient was observed between Mach .2 and Mach .3. The decrease in drag coefficient cannot be readily explained from the data that was collected. Theoretically, the drag should vary as the square of the velocity. The large number of data points taken at Mach .2 and Mach .3 does confirm that the decrease in spinner drag coefflcient is a real phenomenon.

7.3.3.2 Centerbody Drag As mentioned earlier, the Prop-Fan was tested in the presence of a centerbody which was designed to alleviate compressibility losses in the blade root sections. With the force measurement as shown in Figure 7-13, it has been shown that the Prop-Fan net thrust (TNET) cannot be directly measured on the force balance (Reference 21). This Is true because, as discussed in References 21 and 22, the thrust of the Prop-Fan blades changes the pressure acting on the centerbody, thereby changing the pressure drag. The presence of the body also causes an increase in thrust on the rotor equal in magnitude to the change In the centerbody drag. The change In centerbody drag is commonly referred to as the buoyancy force (BF). The measured Prop-Fan thrust has been classically referred to as apparent thrust (TApp) and is the largest force component sensed by the balance. Since the increase in centerbody drag negates the increase In Prop-Fan thrust, there Is no net increase In thrust produced by the Prop-Fan system. However, the thrust measurement (Fs), In Figure 7-13 (Views A and B), does not sense centerbody drag, therefore, measured or apparent trust (TApp) iS corrected by subtracting the buoyancy force, which Is the difference between the centerbody drag measured at the Prop-Fan operating point of interest (DcB) and the centerbody drag measured at the same Mach number without blades (DcBT), The centerbody pressure drag (DcBT), was determined by pressure integration of the longitudinal rows of area-weighted static pressure taps. The integratlon requires a simple summation because the pressures are measured in the centers of equal annular areas.

DCBT " X (PN - Po)AN + E(PM - Po)AM N=I WHERE: Po = FREE STREAM STATIC PRESSURE PN, PN " STATIC PRESSURE AT TAP N, M AN : INCREMENTAL FORWARD CENTERBODY AREA AT TAP N AM - INCREMENTAL AFT CENTERBODY AREA AT TAP M 7.3.3.2 (Continued) As a result of several last minute profile modifications, by ONERA, to the aft end of the Prop-Fan test rig centerbody, the local static pressure taps previously requested by Hamilton Standard were omitted. Therefore, the buoyancy force correction applied to the apparent thrust is In error.

However, the change In the buoyancy force that would have occurred is small enough to result in no significant effect to the data. Therefore, the equation above is further simplified to the following: DCBT = _ (PN - Po)AN N:I Figure 7-13 presents the variation of centerbody drag coefficient without blades (CcBoT) with corrected Mach number (McoR)- The centerbody drag coefficient without blades Is given below and is expressed in terms of centerbody drag force without blades (DcsT), dynamic pressure (qo) and projected centerbody area (Acs).

DCBT CCBDT qo AcB AcB = .22 m 2 A quadratic equation was fitted to the centerbody drag, as shown below and on Figure 7-15 and was used to compute the buoyancy force correctlon to measured or apparent thrust during Prop-Fan performance testing.

CcsoT = 7.6393M" - 16.786M 3 + 12.124M 2 - 3.7165M - .2320 COR COR COR COR It is observed from the data that the centerbody drag coefficient decreases with increasing Mach number. This indicated that the centerbody surface Mach numbers are Increaslng at a faster rate than the free stream velocity.

7,3.3.3 Performance Corrections With the Prop-Fan blades installed and thrusting, as depicted In Figure 7-13, the balance measures the algebraic sum of the apparent thrust, the spinner drag, and the back pressure force. Therefore, the apparent thrust of the Prop-Fan Is obtalned as shown In the foIlowlng equation: T, pp = Fa - F.p + Ds + FT, z

/

r_ W r_ Z U o

./

< O N M w N U I,I n,, U F,, IE O U I i- Z 0 [d i _u

./

IE c_ N O r_ ÷ u n,. _D o < _u n- tO > O O m I n, n,, I,d ,,¢, i- Z -IE Ld U o u_ _K H I-- ILl Q m U U _D m.

? ?

?

I- a m U U 7.3.3.3 (Continued) As defined earlier, the centerbody drag is obtalned from centerbody surface pressure integrations" DCB = E (PN - Po)AN N:!

Also, as mentioned previously, the buoyancy force was obtalned from the difference between these and the tare run pressure Integratlons" BF = DcB - DCBT F|nally, the net thrust (TNET), which is defined as the propulsive force of the blades operating In the presence of the spinner and centerbody flow field wlthout the increase in thrust due to the mutual interaction, is obtained by subtracting the buoyancy force from the apparent thrust: TNET = TApp - BF 69/70 8.0 BLADE STRUCTURAL DYNAMIC EVALUATION 8.1 Test Objectives 8.1.1 Confirm that the SR-TL Prop-Fan was free of high speed blade flutter over the portion of its operating envelope that could be run in the ONERA SI-MA wind tunnel.

8.1.2 Evaluate the IP blade strain sensitivity of the SR-7L Prop-Fan for a range of blade angles, rotational speeds, Mach Numbers and inflow angles.

8.1.3 Compare the measured and analytically predicted IP blade strain responses of the SR-TL Prop-Fan for selected operating conditions.

8.2 Test Procedure Testing of the SR-7L Large Scale Advanced Prop-Fan was conducted In the ONERA Sl atmospheric wlnd tunnel in Modane, France. The Prop-Fan was mounted so that the rotor plane was located in the throat of the wind tunnel. The tunnel throat was eight meters in diameter.

The drive system as described earlier in section 6.2, was supported by rods and cables attached to the tunnel walls as illustrated In Figures 8-I and 8-2.

A balance and torquemeter, also described in section 6.2, provided the capability to measure forces and moments acting on the Prop-Fan.

Test rig vibration was monitored by two sets of horizontal and vertical accelerometers. The accelerometers were located on the drive train housing in two planes aft of the Prop-Fan.

A stationary aerodynamlc falrlng was located downstream of the Prop-Fan. The fairlng served as an extension of the external aerodynamic contour of the splnner. The fairing resulted in an approximate 35% blockage of the flow through the Prop-Fan rotor.

Power available from the test drive system was significantly ]ower than the rated power of the Prop-Fan. Therefore, In order to simulate operation at high power loading conditions, the Prop-Fan was run in two and four blade configuratlons as well as with eight blades. Thls allowed power loadings per blade to be achieved, which correspond to intermediate and high power operating points with elght blades. The disadvantage of operation with two and four blades was the negation of the Inter-blade cascade effects which are present in the eight blade design. These effects tend to be destabilizing in that they lower the Mach Number at which the onset of classical flutter occurs.

The mlsslng blades were replaced with stubs in the two and four blade conflguratlon as depicted in Figure 7-9. The ends of the stubs were machlned to match the external contour of the spinner. The Prop-Fan is shown in the two, four and elght blade conflguratlons in Figures 8-3, 8-4 and 8-5.

?2 \ tn r_ == .J Z u_ Z o Z _3 TWO BLADE TEST CONFIGURAI"IOIM FIGURE 8-3.

ORIGINAL PAGE BLACK AND WHITE PHOTOGRAPH FIGURE 8-4. FOUR BLADE TEST CONFIGURATION Bi_ACK #,_,;u WHITE PHOTOGRAPII FIGURE 8-5. EIGHT BLADE TEST CONFIGURATION ?6 B._'_._ /_,--_u li_,. L PHOT_O_RAPI-f

8.2 (Continued)

The strain gage arrangements for the two, four and eight blade configurations are shown in Figures 8-6, 8-7 and 8-8. The blade surface gage locations were chosen either to correspond with the points of maximum strain for the blade normal modes or to provide the distribution of strain along the entire span of the blade. Blade shank gages were also used to measure vibratory bending moments.

The Prop-Fan was operated in a Beta control mode during the high speed wind tunnel testing. In this mode the blade pltch angle was selected by the operator uslng a manual control. The test procedure consisted of starting the Prop-Fan at a blade angle (B3/4) of 20 ° to 25 ° and increaslng power to obtain 1200 RPM to 1500 RPM rotor speed with the tunnel flow drive system not operating. Mach number was then increased in increments. Foliowlng each increase In Mach number the blade angle was Increased to maintaln the rotor speed In the 1200 RPM to 1500 RPM range. At the Mach numbers of interest, test points were run at two or three different power settings and over a range of rotor speeds. Blade straln gage data was recorded for thirty seconds at each test point. Aerodynamic performance and test rig vibration data were also logged concurrently with the strain gage data.

8.3 Discussion and Results 8.3.1 General Discussion During initial balance runs of the Prop-Fan, which was conducted at zero Math number, test rlg critical speeds were discovered at 360 RPM and 540 RPM.

Both crlticals were highly undamped, allowing rig vibration to grow rapidly if operation was attempted at these speeds. The critical speed at 360 RPM corresponded to the predicted rig first critical. Since both of these crlticals were well below the planned test operating speeds, they did not pose a problem. No other critical speeds were apparent within the test rotational speed range. For the low rotational speeds at which dynamic balancing was accomplished, dynamometer vibratory stresses were the limitlng factor rather than rlg vibrations measured by the accelerometers. This may have resulted from relatively low accelerations causing large displacements due to the low frequency of the response.

Figures 8-9, 8-10 and 8-II present a mapping of the test points run during structural dynamic testlng. The test points acquired for the four blade configuration spanned the entire planned operating range. Power supplied to the Prop-Fan by the turbines was limited to 800 KW due to the elevated ambient temperature in the tunnel.

Operation between lO00 RPM and 1500 RPM resulted in high rig vibration for the two blade and the eight blade configurations. Therefore, this operating region had to be avoided. The vibration frequency was not IP and was not believed to be caused by the Prop-Fan. With the exception of the 2 bladed configuration at 3 ° inflow angle (_), operation at inflow angles of 3° or lO: also resulted In a high level, low frequency vibration which precluded testing at those conditions.

®

FORWARD TO AFT VIEW 3F GAGE r/R TIP ! .44 2 .77 3 .64 4 .69 5 .71 6 .84 7 .71 8 .84 9 .34 I0 .44 I 1 .44 12 .57 13 .78 BLADE CAMBER SIDE VIEW

®

GAGE NUMBERING CONVENTION - X XX FIGURE 8-6. STRAIN GAGE ARRANGEMENT FOR TWO BLADE CONFIGURATION

®

I FORWARD TO AFT VIEW

®

5F GAGE r/R TIP ! .44 Z ,77 3 .84 BLADE CAMBER SIDE VIEW 4 .69 5 .71 6 .84 7 .71 8 .84 9 .34 10 .44 I I .44 12 .57 I 3 .78 GAGE NUMBERING CONVENTION -X XX STRAIN GAGE ARRANGEMENT FOR FOUR BLADE CONFIGURATION FIGURE 8-7.

®

!

FORWARD TO AFT VIEW

® /

®

o,,; BLADE CAMBER SIDE VIEW Z .77 3 .84 4 .69 5 .7!

6 .84 7 .71 8 .84 g ,34 10 .44 1 1 .44 12 .57 ! 3 .78 GAGE NUMBERING CONVENTION -X XX B-A°E.o!.

GAGE NO/ FIGURE 8-8. STRAIN GAGE ARRANGEMENT FOR EIGHT BLADE CONFIGURATION 8O SR-7L 2 NRT TEST POINTS 8OO + SO0 × + Z_ qoo 20O 1800 2000 12O0 I qO0 t600 L06O 8OO RPH SR-'7L 2-HI::I'I" TEST DFITR(HODFINE] ; cp O ( 25OO

po

O ic i z 2000 o !

t 500 i

i

t_ f 1000 800 1000 t 200 I q O0 t 600 1800 2000 RPH FIGURE 8-9, TEST POINTS AND BLADE SHANK MOMENTS (2 BLADE CONFIGURATION) SR-'/L q WRY TEST POINTS I000 t I 1 I P"l MN= .2-.35 (_ DEG. INFLOW C_ MN=.4-.65 {_ D£G. INFLOW) z_ MN= .7-.836 (_ DE:G. INFLOW} r BOO

I

I

f

I 80O == w O

,t, . t

.= m Q.

qO0 B 20O p- .-..v,.j l_iO0 1800 200' 8OO 1200 1000 1$00 RPN SR-7L q-_Rr TEST DRTR {NOORhrE_ I I i / l I 1 _ IRP-GAG£ IF (_ D£G. INFLOW) 2 (3 ]RP-GAG£ 3F (_ DEG. INFLOW) 3 A IRP..GAGE SF (S_DEG. INFLOW) 25OO l _= 1] ,1 rl (9 m A Ii.

th I)

° !

) i

i il

........ J_ . ........... i ............

0 ..... | .... i ....

1000 000 1200 lqO0 1600 1800 2000 RPM FIGURE 8-10, TEST POINTS AND BLADE SHANK MOMENTS (4 BLADE CONFIGURATION) SR-TL 8 NR¥ TEST POINTS I000 I I I I I

I

[] MN: .19-.35 (_ DEG. INFLOW) (_ MN= .4-.65 (_ DEG. INFLOW) O I ) I _,N= .7 i .. (_, =)_G.j INFLOW) 8OO

I

I ot

T

6O0 v I m.

2OO aaoo 800 1000 1200 Iq00 tS00 RPH SR-'7L 8-NRY TEST DATA (HODRNE) 3000 J i _eF._GE: 3F'(_ ocG. INFLOW) I 2 O lAP-GAGE: 4F (_ DI[G. INFLOW) 3 A [I RP.-GAGE:I SFi(_ DE:?. INFLOW)I i 15oo m iooo

o

VA.

!'

50O

]

..... i .......

.... i ........

600 1800 2000 1200 lqO0 800 1000 RPIq FIGURE 8"1 1. TEST POINTS AND BLADE SHANK MOMENTS (8 BLADE CONFIGURATION) 8.3.1 (Contlnued) Mach numbers above .73 resulted in negative thrust for the eight blade configuration due to the low power available from the turbines. Since the balance was not designed for negative thrust, the maximum Mach number for the eight blade case was limited.

8.3.2 Data Reduction The strain gage signals were recorded on magnetic tape for each of the gages shown In Figures 8-6, 8-7 and 8-8. These signals were then fed through a peak stress (strain) converter, a device which determines the peak vibratory strain occurring in a unlt of time (.l seconds for thls data). The output of the peak stress (strain) converter is then statistically analyzed to obtain the average amplitude (mean) and standard deviation of each signal. The IRP (infrequently repeating peak) Is defined, for the purposes of this report, as the mean plus twice the standard deviation. It Is a conservative measure of the strain amplitude normally used to estimate blade fatigue life.

Figures 8-9, 8-10 and 8-II show summaries of the conditions analyzed along with plots of the IRP shank moments for the 2, 4, and 8 bladed configurations.

It can be seen that the highest values recorded were for the two bladed configuration wlth angular inflow (3 degrees). It Is also noted that there are significant blade-to-blade dlfferences. However, it was found that the data contained significant high frequency noise (several thousand HZ), which tended to artificially increase the IRP values. In general, no significant differences In strain gage signal were revealed between the two, four and eight blade configurations at uniform inflow angle, for the same gage (3F).

Figure 8-12 shows an example of the signal to noise problem and its variability between different gages. Because of the low response levels, the data reduction was not repeated with the noise filtered out.

Tables 8-I, 8-11, and 8-111 show the conditions of the runs made along with some sample IRP tabulations, showing blade-to-blade variations.

Table 8-1V shows a tabulation of peak IRP values for each gage versus the limits determined based on the threshold at which blade fatigue damage begins to accumulate. Because of the high noise content, all of the IRP values (Figures 8-9, 8-10, 8-II and Tables 8-I, B-If, 8-IIl and B-IV should be interpreted as conservative upper bounds on the strains actually felt by the blades. As such, the gages shown to be exceeding the Iimits should actually be Interpreted as below the limits with the noise filtered out.

W .J (.9 Z < ..J Z 14= :E 0 Z < F- Iv o < C3 1-1 In J l,,IJ LO I < z LO W U) l-- U Z Ld ILl z Q < ,..I v Z < "r -i Z I.I.I (/I < P < ,J -I 14.

n_ o I.I.

LI.I (/I 0'11_ z Zo Z o I- ,..I < < _:e4 < Z I1_ II (3 uJ Z (/I -I, zlE I,i.

O_ Z w 0 w n_ z < X W < z U _0 e,i | co 14J n_ (3 I.i.

Z O b- o= m u.

Z O o t_ -I w X O F- -/ I- n_ W -I t_ X A n_ O (/1 X O m !- O o Z l- t_ 0.

O W -J ORIGINAL PAGE I_.

OF POOR QUALIT_

Z

< n_ z L_ w r_ < ..]

m Z _I n n_ ILl B.

r_ Z < A z :E n_ Z i- Z u z I-- ul I!.

W -J II1 !,- A Z V.

Z O o w ORIGINAL PAGE rS 88 O]_ P.OOR QUALITY ORICENAI/ PACE IS OF, POOP, QUALITy Z O Y- < rr O m B.

Z O _9 ILl < -] Ill Z O I-.

< -] m < F- 11.

n, Ill ..l < Q Z < A Z O Z O m Z O (3 (.9 Z 1- < n, I1.

O m !

0D 1.11 --1 m < i-- TABLE 8-]_Z'. MEASURED VIBRATORY PEAKS VS. LIMITS Peak IRP Peak IRP Peak IRP Peak IRP Stain observed Strain observed Strain observed Strain observed for 2- way for 2- way Gage Description for 8-way for 4-way (0' inflow) (O' inflow) (O' inflow) (3' inflow) fraction of limit fraction o! limit fraction of limit fraction of limit Radial Bending (44%r/R) Gage 1 .43 .33 .79 Radial Bending (71%r/R) Gage 2 1.12" .48 1.05" Radial Bending (84%r/R) Gage 3 ,22 .45 .46 .88 Trailing Edge Bending Gage 4 .55 .6 1.01 * T. ,..

Shear (71%r/R) Gage 5 .04 .07 .05 .11 , i Shear (84%r/R) Gage 6 .05 .07 .08 .12 Chordwlae (71%r/R) Gage 7 .22 ,21 .16 .26 Chordwise (84%r/R) Gage 8 .36 .23 .20 .33 Radial Bending (34%r/R) Gage 9 .41 .30 .72 Shear (44%r/R) Gage 10 Chordwise (44%r/R) Gage 11 .24 Radial Bending (57%r/R) Gage 12 .26 .45 ,49 1.11 * .52 1.03" Radial Bending (78%r/R) Gage 13 .29 .29 .76 .28 Flatwise Shank Moment - Gage F .29 .47 ,22 .17 .19 .27 Edgewise Shank Moment - Gage E * It is fell that because of the noise present, these values can be Interpretedas within limits.

Notes: 1. The gage limits were set based on the strains reaching a Ikmslmld at which blade fatigue damage begins to accumulate (vibratory strain _lmrimposed upon steady strain).

2. Becauseof the amount of noise in the strain signals, the Blip values listed here should he interpreted as conservative u_ Immds onthe strains actually felt by the blades.

ORIGINAI_ PAGE IS OE POOR QUALIT_ 9O 8.3.2.1 Angular Inflow As previously noted in section 2.0 and paragraph 8.3.2.1, the only angular inflow data obtained was for the two bladed configuration at 3° inflow angle. A spectral analysis of each strain gage signal was performed using an FFT (Fast Fourier Transform) algorithm. This analysis employed a Hanning window to help overcome leakage effects. Correlatlons were then made with the predictions of IP response.

Table 8-V shows the P-order content obtained for the 3 degree conditions tested (shank moment and radial bending). It can be seen that the 2P component averaged about 25% of the IP component (with 2P/IP approximately .54 near 1500 RPM) and that the 3P component was generally much smaller. It Is noted that at least a portion of the excitation force driving the 2P response may have come from the test rlg drive system. Twice per revolution vibration Is characteristic of a shaft with universal joints. It is also likely that nonlinearities In the aerodynamic excitation, possibly due to observed vortex loading phenomena, or tunnel flow irregularities in the test section, are causes of higher order excitation. The aerodynamic analyses used for the predictions do not consider this effect.

The data digitized for the spectral analysis was sampled at constant time steps. Even small variations in RPM are known to result in substantially lowered P-order magnitudes. By synchronously sampling the data (digitizing at constant fractions of a cycle instead of constant time step) this problem can be overcome. This is known as a 'speed corrected' spectral analysis.

However, since the RPM trace was not available for this test, this was not possible here. An alternative procedure was used for several strain gage results to double check the amplitudes obtained from the non 'speed corrected' spectral analysis. A calculation was made of the average amplitude of a band pass filtered (IP frequency mid band) strain record.

This procedure is cumbersome but should produce conservative strains because the amplitudes are those generated by a peak stress converter (with a finite reset rate). Table 8-VI shows that the band pass filtered values are typically I0% higher than the spectral values. Even with no RPM variation, the spectral magnitudes are known to be up to 16% low, due to the effect known as leakage (with a Hannlng window as employed here) In digital signal processing terminology. It is concluded that the actual magnitudes are within lot (higher) of the spectral values.

Figures B-13 and B-14 show the trends of IP response (shank moment and blade bending) with respect to power, RPM and Math number for the 2 bladed configuration at 3° inflow angle. As expected, the response increased with power and Mach number. Variations of RPM are influenced by system resonances. Figure B-15 shows the variation of bending strain with spanwlse location. Again, the trends are believable and consistent with the data collected during the four and eight blade structural testing.

ORIGINAl5 PAGE Ig OE POOR QUALITY i+ + + + ÷ + + i0 o,0 0 0 0 mLK_ 010 0:0 0 0 0 0 0 0 0 ¢1, 0 O 0 0 0 0 O. 0 _ ¢_.

''_ o o o o o o o o olo o olo I q _!_ _ _,a _,a o .., o _ .... o ........

• . ......... _ " ¢/ _.

J • . , ........ -- .

Z i !- < I- • _,_.:_i_,_.: ....... _,_,_:,s._ .... _!_ _ -,_ ....

< Z u < l- u.

I- Z /,i I- z U r_ td o e_ I!.

Ill .J < I- o o o. o o.

z0_,,_ _.

o.

o o

0 _ : _._--

o o o IP III _,' _.0 _.- ._ _ ._ _ .: _o .g.

¢_ _ "_- o. . .

° |

_i m ,_ _ : o. o. o. o o.

C C C E £ v A II :E v >- I.= l- Z 14J p - _ _E z w in o \ o i- Z M) rv < IU -r _X (n .,0 O 0¥ O Lz O -I o !.1.

O= O Z 1. tn td n, <

iiiii

-I ILl .I-I- L9 _0 O Z _. m O < 0 O 0 O 0 o O O O 0 0 0 0 r_ _ m ILl Z> (..-13/i,M-N) AJ.IAI.LISN3S .I.N_II_IOIAI iiiii Q _P • INVHS 3SIMJ.Y"] 4 d!

< IIIII < .J 0II _u N O o < O_- O lllIl 6'1 In A • i!

l_!il Z O IiiiI - v / !i!!!

h, _= i :, _ .. ",, __ _Z m z n, 0 0 0 0 0 0 / o o 0 0 0 0 0 0 Ig) / / Q Q [d z Z (_3/_°N) X.LIAI.LISN3S .Z.N3WOW _NYHS ]S|MJ.V'IJ d| _.J O I- < IEIE z _4 )!0 _,< II0 0"1" I- W l. ln n, ul Ill o O O bl l, o o _0 O O O O O o o 00 O O O O O o (_13/l_l-N) XJ.IAI.LISN_IS J.N31_Oi_I _INVHS 3SIMJ.V'I_i d i !

Z w i "1 Z z LU --I <

B

< n, r i w \\ ",..,.

-I I, Z

\.!._

n, < I!!!!

-I (.9 Z < ..'. iiiii (3]/N|BMLSO_31N! IL]AiLI_N]$ N[UVLS dl C] I1|11 < ..I m _ -ttt|i z 111111 N: (n

I

Z !

t- J.

\r :, Ii.

z

/it

r,.

// .." / _LIAILISN3S NIUHIS d!

l-

!

Y ,i 1t[ "r i | |111 U, II|#IIItAI||OM_IW LLIAILIINH N1¥¥[| dl 9G SRTL=IP BENOING STRQIN SENSITIVITY VS. SPRN(MN=.S.5OOKW} i _ LSOO I BOO T" ;' ,I ,'7 lqO I z I w

! ]

lEO

i i

i !

OO i i !

BO I l

f

r

t " t g _0

I

z

I J

!

w I i i I i ....

L l,O000 o )._ooo FRRCTION OF SPRN SRTL=IP BENDING STRRIN SENSITIVITY VS. SPRNtMN=.TS.570KN} 160 ,.

I ..... J 700 - '- IOOO 'M ,T lqO z g LI I00 "%% x !

\ , > BO l-,- °_.°. - • ".,[-A - • _ : :z o uJ

o,. i

t o 6 " " o._o6- - o;.oGo " oZeo_o - =.oooo FRRCTION OF SPRN FIGURE 8-15. RADIAL BENDING VS SPAN, 2 BLADES, 3° ANGULAR INFLOW 8.3.2.1 (Continued) The plotted values were normalized wlth respect to E.F. (excitation factor) which Is defined as: E.F. = _ (VT1644.8)2(p/po) Where _ is the inflow angle In degrees, VT the true airspeed In Km/HR, p the alr density, and po the alr density at standard sea level conditions.

Only the IP component of the response has been plotted to investigate trends.

The IRP values include the higher order content as well as unfiltered noise.

8.3.2.2 Analytical Method A finite element model of the SR-7L blade was developed during the design phase. The same model was used here to predict IP response for correlation with test. The steady state aerodynamic loads were calculated using the HS computer code H444. The steady alr loads along wlth centrifugal loads, were applied to the model which was analyzed using the in-house finite element code, BESTRAN (Reference 23). From this solution the differentlal stiffness matrix was obtained and added to the structural stiffness matrix. The dlfferential stiffness matrix represents the addltlonal stiffness of the blade under centrlfugal loading. The unsteady (IP) alrloads were then calculated uslng the HS computer code H337. The loads were applied to the finite element model and displacements and surface strains predicted.

Mohr's circle relationships were employed to calculate strains In the direction of the gages for correlation with test. Shank moments were calculated us|ng the reaction loads at the root of the blade and assuming a linear variation of moment up to the shank location.

The H444 and H337 codes are both aerodynamic 'strip' analyses that have been calibrated to predict the steady and unsteady alrloads on swept Prop-Fan blades. The H444 code employs a Goldsteln formulatlon to calculate induced velocltles, whereas the H337 code uses a skewed wake theory. Two-dlmenslonal, compressible alrfoll data was used by both codes to predict the aerodynamic loads. Trlangular plate elements were used In the BESTRAN code to do the finite element analysls. Each component of the composite blade (i.e., shells, spar, foam, sheath) were represented by separate plate elements. They were a11 tled together using constraints based on plane-sectlons remaining plane.

Further detail on the codes and their use can be found In the LAP Blade Design Report (Reference 14).

Figure 8-16 Illustrates a Campbell diagram of the SR-7L blade. Shown are BESTRAN frequencies compared wlth test values. It is assumed that the RPM range tested was far enough away from IP resonance that the Influence of damping on strain magnitudes can be neglected.

O-FREQUENCIES FROM WPAFB TEST I-PRE-TEST PREDICTION (BESTRAN) ZS0 ZOO N 'r ).

u Z b4 bJ I_0 -- ml < ¢ I" <

oz f. v

o- I I I o soo _ooo isoo zooo zsoo ROTATIONAL SPEED"" RPM FIGURE 8-16 COMPARISON OF SR-7L NATURAL FREQUENCY TEST RESULTS TO PREDICTIONS 9B @.3.2.3 Test vs. Analysis Three cases were chosen to analyze and compare to test. The condltions were picked to study the influence of power and Mach number variations on the predicted strains. Case l (run #I176) was at 1673 RPM, 794 KH, .5 MN.

Case 2 (run #1190) was at 1698 RPM, 244 KN, .5 MN. Case 3 (run #1214) was at 1707 RPM, 829 KW, .724 MN.

The IP gage strains were calculated using the methods outllned in the prevlous section. Comparisons were then made to the values obtained from a spectral analysls of the test data. See Figures 8-6, 8-7 and 8-8 for gage locations. Figure 8-17 shows the variation of radial bending straln as a function of spanwise 1ocatlon. As can be seen the correlation is quite reasonable wlth the In-board predicted strains being about 30% too high.

Nowever, as previously noted in paragraph 8.3.2.1, the 'measured' strain levels are felt to be up to I0% low.

Figure 8-18 shows the correlatlon of shank moments and root radial bending (gages F and l) as a function of power, (holding RPM and Mach number). Both the test and analysls show the expected trend of increased IP response wlth Increased power. The test values are consistently below those calculated.

Figure 8-19 shows the same quantities plotted against Mach number (holding RPM and power). The trends are consistent with expectations (higher response with increased Mach number). Table 8-VII shows a comparison of all the strain gage results for the three conditions. As noted from the plots, the correlation with radial bending strains and shank moments Is quite reasonable. However, the correlation Is rather poor with the gage placed near the trailing edge, and also with the chordwlse and shear gages. This Is consistent wlth previous experience. It |s felt that these values are affected more by local distributions of aerodynamic loads, whereas the radial gages are strained by an integration of loads outboard of a given gage.

Further study Is needed to better define the local load distributions In order to more closely correlate wlth the measured strains.

8.3.2.4 Blade Flutter Prior to the high speed wlnd tunnel test, an unstalled flutter analysis was conducted for the two, four and eight bladed Prop-Fan configurations. The analysis indicated that the Prop-Fan would be free from unstalled flutter throughout the wlnd tunnel test operating envelope. The tests confirmed the predictions In that no flutter or tendency to flutter was measured or observed. The wlnd tunnel test results were also a good Indlcator that no unstalled flutter would occur durlng the subsequent flight test program.

Though the maximum Mach number achieved for the elght bladed configuration was 11mlted to .73, the resultlng dynamic pressure at the tunnel's effective pressure altltude of 4,360 meters (14,OOO ft.) Is greater than the maximum dynamic pressure expected for the f11ght test.

|

J

oJ

.=.

/

z:

/

(3 Z < n

i L °

> !

Z c /m _o -~ e.

Z

/

III -I o ,,-- / < i o Q z < n, u) m ul >- ..I < Z < (n > E p.

_ -,,, bJ !-- | l | a "7 co z

/

bJ

/

m

0_

ru | J Jl | NI_LS;OV3lll '1¥10H¥ 4!

lOO T_'STVS. ANALYSIS-- EFFECT OF POWER (AT 1700 RPM;MACH NO. =.5) 1 _ TIST _FLATWI ;E(GAGE 3F} 2 _) _IALYSISiFLRTWI'iE(GAGE F I 3 ""_'" ST TEDGEWI _E {GAGE 3E} Z000 U _- AIAI Y_TSIFn_FWT ;F (C,AC, F F :E i-, 1500 !

Z &d 3E RUN # C 1176 X lOOO !

z RUN # 0 , _.. I Z L

I '

L. i

0 ......... _:':'"':"':",""L": "''°'" -'_':2"L ................... _ . . . { ....

o zoo 400 600 8oo iooo POWER (KW) TEST VS. ANALYSIS-- EFFECT OF POWER (AT 1700 RPM; MACH NO, : .5) 1 _ TEST -AVG. F 31A_O 7] I 2 C ANAL'SIS- G GE 1 5OO Z < n, J- m 0 300 U C RUN # m a.

RUN # m ZOO 1190 _ _'Z IO0 t ....

400 600 800 1000 0 200 POWER (KW) FIGURE 8-18. TEST VS. ANALYSIS (EFFECT OF POWER) lOl TEST VS. ANALYSIS-- EFFECT OF MACH NO. (AT 800 KW; 1700 RPM) 2 O ALTSI5 FLATWI IE(GAGE F | o RUN1214 # 3 - -,_,.-. ST "DGEHI _E (GAGE __L _-J- AI Y_IS ENr,FWT _Ff_,Ar, F F 2O00 :S J .,.

l- Z 1500 ld :E RUN # J !176 10o0 Z < Z a.

b .......... ! ...........

0.9000 0.4000 0.5000 0.6000 0.7000 0.8000 MACH NO.

TEST VS. ANALYSIS-- EFFECT OF MACH NO. (AT 800 KW; 1700 RPM) I _ TIIST-RVG OF 31 RND 71 2 O QIIRLYSIS GRGE RUN # O 1214 Z 4O0 m < r,, I- u) O f u RI_IN # / !176 :i a.

IOO

!

0.4000 0.5000 0.6000 0.7000 0.8000 0.9000 MACH NO.

FIGURE 8-19. TEST VS. ANALYSIS (EFFECT OF MACH NO.)

ii ,i

E ,: ..: _ .: - ,.: _ 0, : .'_ li.

,z,=,, ill

i|_ '"

m ,jg e- (il >- ,.,I < Z < > l I-, I.Ll I-' 1.1..

_i ° ""

,< O,.

I,Ll -1 lid <

_1_-,_ o o o "-,!

I- i ,.:e_ _ _ _:

i_; " li #fiiii; "

ill l_, lit lit.i lil

i 8.3.2.4 (Continued) The maximum Mach numbers achieved for the two and four bladed Prop-Fan configurations were .80 and .84, respectively. Though these configurations did not permit the effects of blade cascade to be evaluated, they did demonstrate the Prop-Fan's resistance to unstalled flutter onset for subsequent blade surface pressure testing.

8.3.2.5 Whirl Flutter Due to the flexibility of the ONERA test rig, whirl flutter was a concern for this test. These concerns were reinforced when initial calculations of the installations' stability showed whlrI flutter to occur within the operating envelope. For the Inltlai computations, the forward support structure, shown In vlew A of Figure 8-I and views A and B in Figure 8-2, had a three rod configuration that resulted In undesireable symetrical pitch and yaw stiffness. To stabilize the two and four bladed configurations, a two rod support structure, as depicted In vlew A of Figure 8-I, was utilized. The two rod configuration, which provided asymetrlcal pitch and yaw stiffness, resulted In a more stabilized system. The computations also demonstrated the need to add cables, as shown In view A of Figure 8-2, for the eight bladed configuration to increase structural damping while improving the distribution of the pitch and yaw stiffness. The end result of these rig modifications was successful completion of the wind tunnel test with no whirl flutter instabilities encountered.

9.0 AERODYNAMIC PERFORMANCE EVALUATION 9.1 Test Objective To measure the aerodynamic performance of the SR-7L Prop-Fan for a range of blade ang]es, rotational speeds and Math numbers.

9.2 Test Procedure Aerodynamic performance data (thrust and power) was measured concurrently with the structural dynamic data. As a result, the data Includes losses In performance resulting from the presence of the strain gages and strain gage wires on the aerodynamic surfaces of the blade. Installation of the gages was accomplished so as to minimize these losses.

Power avallable from the two turbine engines driving the Prop-Fan in the wlnd tunnel was slgnlficantly lower than the rated power of the Prop-Fan.

Therefore, testing was conducted using two and' four blade configurations as well as eight blades.. The two and four blade configurations permitted operation at power loadlngs per blade that correspond to high and Intermediate power operating points respectlvely for the eight blade Prop-Fan deslgn. The disadvantage to thls approach is that the eight blade performance cannot be easily extrapolated from the two or four blade test results. This is due to the aerodynamic Interact|on between the blades. This interaction is more significant for the eight blade design than for the two or four blade conflguratlons, due to the reduced spacing between blades. Unfortunately the aerodynamic performance data for the 2 blade configuration was llmlted, and rather than attempt to project the full power elght blade performance from the the two and four blade test results, the available results are compared to analytical predictions. This serves to verify the analytical techniques and provides confidence that the predictions for the eight blade high power performance may also be correct.

The missing blades were replaced with stubs (see Figure 7-9) in the two and four blade configurations. The ends of the stubs were machlned to match the external contour of the spinner.

The Prop-Fan was operated in the Beta Control mode during the aerodynamlc performance testing. In this mode, Hamilton Standard personnel were able to change the blade pitch angle during testing by means of an Increase/decrease pltch switch located in the control room. For a fixed Mach number and a constant power supplied by the turbines, the Prop-Fan rotational speed was varied by increasing or decreasing blade pitch angle. At the Mach numbers of interest, aerodynamic performance data was collected for two or three dlfferent power settings and over a range of rotational speeds.

Approx|mately 140 performance data points were collected.

9.3 Discussion and Results

9.3.] Data Reduction Procedure The blade pitch angle (B314) was measured and recorded for each test point.

An electrical signal proportional to blade angle was provided by means of a potentiometer mounted on the rotating side of the Prop-Fan. The potentiometer shaft was positioned by the rotation of the No. 7 blade retalning ring through a cable and pulley arrangement (Figure 5-3). Flgure 9-I demonstrates the linearlty of the output of the blade angle measurement potentiometer as monitored during a pre-test calibration check.

The Prop-Fan rotational speed was measured by use of a IP pickup. The sensor was triggered by a gear _K)unted on the test rig drive shaft. The rotational speed was averaged over ten revolutions.

The power absorbed by the Prop-Fan (BHP) was determined by multiplying the torque supplied to the Prop-Fan (QcoR) by the rotatlonal speed (N). Torque supplied to the Prop-Fan (QcoR) was computed by subtracting the measured frictional losses in the balance (Q_L) from the torque measured by the torquemeter (QMEAS).

BHP = N x QcoR Where: OcoR - OM_s - Q_L The net trust (QNET) determlned during testing is the unlnstalled thrust of the Prop-Fan rotor, operatlng in the presence of a spinner and centerbody and Is computed from the following equation" QNET = F8 + Ft. - Fs, + Ds - BF Where" F, = axial force measured by the balance FTH = losses due to thermal effects F,p = back pressure force Ds = spinner drag BF - buoyancy force The temperature correction term (FTH) compensates for the effects of changes In temperature on the balance strain gages.

The back pressure term (FBp) corrects for the increase In measured thrust due to the differential between the pressure behind the spinner bulkhead and the free stream pressure. The back pressure force Is calculated by multlplylng the difference between the average pressure measured by the taps shown In Figure 7-I0 and the free steam pressure by the pro_ected area of the spinner bulkhead.

100.

7_ BLAD£ ANGL£ (_3/4) I !000. 1500. 2000.

-500. O. 500.

-1500. -I000.

-ZOO0.

PO'rENTIOMETER OUTPUT (mV) FIGURE 9-1. LAP BLADE ANGLE POTENTIOMETER OUTPUT 9.3.1 (Continued) The splnner drag force (Ds) is computed by multiplying the spinner drag coefficient by the free stream dynamic pressure and the reference area of the spinner, as defined in paragraph 7.3.3.1. The spinner drag coefficient was determined as a function of Mach number during the spinner drag test (Figure 7-14).

The buoyancy force term (BF) eliminates the apparent Increase in thrust caused by the interaction between the Prop-Fan rotor and the centerbody. The buoyancy force is determined by measuring the centerbody drag at the performance test point of interest and subtracting the centerbody drag without blades for the same Mach number. The centerbody drag was determined during performance testing by integrating the difference between the pressures measured by the taps shown in Figure 7-I0 and the free stream static pressure over the surface of the centerbody. (See paragraph 7.3.3.2). The centerbody drag without blades was computed by multlplylng the centerbody drag coeffi- c|ent by the free stream dynamic pressure and the centerbody reference area.

The centerbody drag coefficlent was determined as a function of Mach number during the splnner drag test as depicted in Figure 7-15.

Mach number was established from the ratio of static pressure (measured four meters upstream of the Prop-Fan rotor) to stagnation pressure. Static pressure was also measured In the plane of rotation as a backup. The ratio of static-to-stagnation pressure was correlated with data taken during a pre-test ca]ibratlon in order to compute the Mach number (see Paragraph 7.2).

The Prandlt-Young correction was applied to the computed Mach number to compensate for the effects of the tunnel walls and the thrust produced by the Prop-Fan.

9.3.2 Data Presentation The most complete aerodynamic performance data was acquired for the four blade Prop-Fan configuration. Operational problems encountered with the test rig, while running the two and eight blade configuration, limited the operating envelope for these configurations. The r|g operational problems were addressed earlier In section 8.0. The structural failure of the centerbody during the test also significantly delayed the test program. Therefore, In order to expedite the program, the test points were limited to the boundaries and a few Interior points of the operating envelopes for the two and eight blade configurations.

9.3.2 (Continued) The aerodynamic performance data was nondimenslonalized for analysis according to the following set of equations.

(POWER COEFFICIENT) BHP (po/p) Cp= 5.674 (ND/IOOO)3D 2 (NET THRUST COEFFICIENT) TNET(Po/p) CTNET = 340.42 (ND/IOOO)2D 2 V (ADVANCE RATIO) J = 60 -- ND Where: BHP = power, KW TNET = net thrust, N D = Prop-Fan diameter, m po/p = density ratio, sea level to ambient N = rotational speed, RPM V = free stream velocity, m/sec It should be emphaslzed that the performance data was acquired only during structural testing where blade angle was continually varied to maintain a constant power level. Accordingly, the data was plotted as curves of power coeffIclent and net thrust coefficient versus advance ratio to eIimlnate blade angle (63/4) as a variable. If desired, thls data can be converted to efficiency (n) by the relationship: (EFFICIENCY) CTNeT q= X,,1 Cp The power and thrust coefficlent data as a function of advance ratio are presented for the four blade conflguratlon in Figures 9-2, 9-3 and 9-4 and for the two blade conflguratlon In Figure 9-5. The data was then cross-plotted to derive the more conventional maps of CTNET versus Cp.

Plots of CTNET versus Cp for the two, four and eight blade cases are presented In Figures 9-6, 9-7, 9-8, 9-9 and 9-I0. The llmlted scope of the two blade and eight blade data Is apparent. Data taken at 3 ° inflow angle rather than at 0 ° Is presented for the two blade conf!guratlon, because a better dlstrlbutlon of test points was run at that angle. Examlnatlon of these plots shows that the net thrust coefficient exhlblts smooth consistent varlatlons wlth power coefficient and advance ratlo.

SR-7L WIND TUNNEL

PERFORMANCE DATA

• O9 f

/

O8 /

/

/ .O7

/

CTNET

/

Cp "" "--"

I P

_Z_ --800KW .o6

/

O =500KW

C] =2!;0KW /

/

/

.05

/

of

/

/

/

/

.04 CTNET

/

.03

If

.O2

1o

J

.01 | • 0 im 0.6 0.8 1,0 !.2 1.4 1.6 ADVANCE RATIO, J FIGURE 9-2. Cp AND CTNET VS J, 4 BLADES, MN = .2, _J = 0 °

SR-7L WIND TUNNEL

PERFORMANCE DATA

.24_ .ZZ 1,0 .Z0 .1 _ .8 .16 .14 .7 CTNET Cp o6 .IZ .5 .10 .O8 .4 m ,3 .06 .Z .04 m .I .02 1.8 2.0 Z.Z 2.4 Z.6 Z.8 3.0 3.2 3.4 3.6 ADVANCE RATIO, J FIGURE 9-3. CpANDCTNETVS. J, 4BLADES, M N=.5,¢=0 °

SR-7L Wi ND TUNN EL

PERFORMANCE DATA

P = 800KW O P = 500KW .10 .!o [] P = 300KW

,,I l

.09 .9 .o8 .8 .o7 .7 CT NET Cp .06 .6

f

.5 .05 .4 .O4 .03 .3 .02 .2 .01 .I 3.2 3.4 3.6 3.8 4.0 4.2 _4 4.6 4.B ADVANCE RATIO, J FIGURE9-4. CpANDCTNETVS. J, 4BLADES, M N=.8,_=0 °

SR-7L WIND TUNNEL

PERFORMANCE DATA

1.0 _ .50

I I f'_

.09 _ .45 CTNET / .

Cp m_

/

A P'- 800KW .08 _ .40 O P'50OKW jJ [] P=250KW _ .07 _ .35 ,. .t / .06 _ .30 CTNET Cp J .05 _ .25 _i 1 i ,_ ,04 -- .20 .03 _ .15 .02 _ .IO

_m" /

.01 _ .05 1.9 2.0 2. I 2.2 2.3 2.4 Z.5 2.6 ADVANCE RATIO, J FIGURE 9-5. CP AND CTNET VS J, 2 BLADES, MN -- .5, _=3 ° m

SR-7L WIND TUNNEL

PERFORMANCE DATA

.32 .30 .28

/

.26 .24 .22 .20 .18 CT NET .16 ,14 .12 ,I0 .08 .06 .04 L_ ,02 ,10 ,20 .30 .40 .50 .60 .70 .80 .!)0 1.0 Cp FIGURE 9-6. CTNET VS Cp, CONSTANT J, 4 BLADES, MN = .2.

SR-7L WIND TUNNEL

PERFORMANCE DATA

.20

/

.18 .16

,//

.14 .12 J CTNET .10

/j//

.08

/,//,'

.06 ___lr J = 3.5 3.0 .04 .J = 2.5 .02 / I J=2,0 • 10 .20 .30 .40 .50 .60 .70 .80 .90 1.0 Cp FIGURE 9-7. CTNETVS Cp, CONSTANT J, 4 BLADES, MN " .5, _=0 O

SR-7L WIND TUNNEL

PERFORMANCE DATA

,16 .14 .12 .1o • CTNET .08 .06 ....

.02 4.4 J=3. =4.0 .60 .70 .80 .10 .20 .30 .40 .50 Cp FIGURE 9-8. CTNET VS Cp, CONSTANT J, 4 BLADES, MN = .8, ¢ =0

SR-7L WIND TUNNEL

PERFORMANCE DATA

.IO J=2.4

/

.o9

/

.o8

J/

.07

"

,o6 CTNET

.05 /

,04 .O3 .02 SJ,, .01 .I .2 ,3 ,4 Cp FIGURE 9-9. CTNET VS Cp, CONSTANT J, 2 BLADES, M N : .5, _ : 3 °

SR-7L WIND TUNNEL

PERFORMANCE DATA

,26-- J= .91 r .:'4- mMN=.t9 mm m,,m, MN =.50 .22- ,20- J=.8 .18- .16-- J=2,64 J 14-- / / CTNET / .12-

/

J=2.31 / / / J °10_ / /7 / .08r

/

//"

/

/

.06- / .04., .02- .5o .60 .10 .20 ,30 .40 Cp FIGURE 9-10. CTNET VS Cp, CONSTANT J, 8 BLADES, _ = 0° 9.3.2 (Continued) The four blade and eight blade data was cross-plotted again to obtain curves of net thrust coefficient versus advance ratio for a constant power coefficient. Comparison of Figures 9-11 and 9-12 shows that for the same advance ratio and power coefficient, a higher net thrust coefficient is obtained with the four blade configuration than with eight blades.

Therefore, as expected, the four blade configuration, at lower power loadings, is more efficient than the eight blade Prop-Fan.

Comparlsons of the calculated and experimentally determ|ned performance of the four and elght blade Prop-Fan configurations are shown In Figures 9-13 and 9-14. These calculations were made using a refined lifting l|ne method.

The predlcted and measured performance agree very well for the four blade conf_guratlon over the entire range of test poInts. Although the performance of the eight blade Prop-Fan design was underpredicted at Mach 0.5, good agreement between measurement and prediction was obtalned at Mach numbers of .70 and .73. Moreover, the trends of thrust w_th power are predicted accurately.

SR-7L WIND TUNNEL

PERFORMANCE DATA

0.28

Cp = 0.10

A 0 Cp = 0.20m

0.24

Cp = 0.40

0.20

[I_ _Cp = 0.30

0.16

CT NET

0.12

ADVANCE RATIO, J FIGURE 9-! I. CTNET VS J, CONSTANT Cp, 4 BLADES, _ = 0 °

SR-7L WIND TUNNEL

PERFORMANCE DATA

0.28

OCp = 0.2

El Cp = 0.24

0.24

Cp = 0.29

o16 _\_

CTNET

0.12

0.08

0.04

5.0

0 1.0 2.0 3.0 4.0

I ADVANCE RATIO, J FIGURE 9-12. CTNET VS. J, CONSTANT Cp, 8 BLADES. _J = 0°

SR-7L WIND TUNNEL

PERFORMANCE DATA

f

/

/

/

/

/

.12

/

.11

/

/

/

/ .IO /

/ I ,//

.09 v

c l /

.o8

i /

I

.07 CT NET

/ 1

/

.06

/ /

CALCULATED

/

/

PERFORMANCE: .05

/ /

M= 0.2, J= 1.0

/

M: 0.5, J= 2.0 "" " "" -- .o4 M---- 0.5, J= 3.0 _=,,m TEST DATA: .O3 I (_ M= 0.2, J= 0.95 r"IM= 0.5, J= 2.0 Z_ /_M= 0.5, J= 3.!

.o2 .01 0 .....

0 .05 .10 .15 .20 .25 .30 .35 .40 .45 .50 Cp FIGURE 9-13. COMPARISON OF MEASURED AND PREDICTED PERFORMANCE, 4 BLADES

SR-7L WIND TUNNEL

PERFORMANCE DATA

c ii

III / ,/

CALCULATED PERFORMANCE: M =.5, J = 2.17 i f M =.5, J = 2.64 t .11 ' .n M =.70, J = 3.0_ .= 73,J:36 / if" / / / .10 / TEST DATA : if r / O M =.5, J = 2. I7 ii I I I / oif ,/, A M =.7, J= 3.04 it .09 ° M =.5, J=2.64 C I /" / I_ M =.73, J= 3.59 / / f .oe / if / i

/ ,/ /

.o, I ,"x' _t

CTNET I,il // I .o5 it

it// I

/ /

.o4 / , I I / .03 .oz .oi .05 ,10 .15 .20 .25 .30 ,35 .40 .45 .50 Cp FIGURE 9-14. COMPARISON OF MEASURED AND PREDICTED PERFORMANCE, 8 BLADES 123/124 lO.O BLADE SURFACE STEADY PRESSURE MEASUREMENT I0.I Test Objective To measure the steady pressure distribution on the surface of the SR-7L Prop-Fan blade for a range of blade angles, rotational speeds, and simulated flight Mach numbers.

10.2 Test Procedure In preparation for blade surface steady pressure testing, the Prop-Fan was installed on the drive system, as described In section 6.2, and supported as shown In Figures lO-I and I0-2.

As noted earller in sectlon 2.0, collection of blade surface steady_pressEre_ test data was interrupted during the first tunnel entry In early 1986 and had to be rescheduled for a second tunnel entry In early 1987. Though the steady pressure test data collected was quite limited as a result of the interruptlon, enough Informatlon was obtained to indicate that a revised pressure tap layout was desirable. A new steady pressure blade incorporating an improved pressure tap layout was fabricated for use during the second tunnel entry. It provided a much higher density of surface pressure taps in areas on the blade where the local steady pressures were more sensitive to changing operating conditions.

The steady pressure measurement blade (S/N 009) was Installed in blade posltlon 7. For balancing purposes, a counterweight blade (SIN 058) was Installed in position number 3, as shown in Figure I0-3. As a precaution, signals from 2 shank mounted straln gages were monitored throughout the test. Special contour matching blade stubs were Installed in the remalning 6 hub arm bores. The test rig power capabllitles necessitated conducting all testing using a 2 bladed Prop-Fan conflguratlon, thereby permittlng operation at power loadlngs per blade corresponding approximately to the take-off and cruise condltlons of the eight blade Prop-Fan design.

The Prop-Fan was operated In the beta control mode during the entire test.

In thls mode, Hamilton Standard personnel were able to change the blade pitch angle during testing by means of an Increase/decrease pitch switch located in the control room. For a fixed Mach number and a constant power supplied by the turbines, the Prop-Fan rotatlonal speed was varled by increasing or decreasing blade pitch angle.

The blade pitch angle (6314) was measured and recorded for each test point.

An electrlcal slgnal proportional to blade angle was provided by means of a potentlometer mounted on the rotatlng side of the Prop-Fan as Illustrated in Figure 5-3.

• ., r,_ /.L-._' ._--_-- _;" _ _-'-----N_ m_ BALANCE GEOMETRICAL CENTER

CABLEB \ A "l zo R

c

"\11"_'_ " I"-I I I _h'--'-'--"--_

CABLES ROD ROD ROD (_: 1 4, 3 L:3700 ...... +r-- t' L::3550 : :: 00 A B L:_,,,o C,

AL- DIME",,ON" IN M,,.'I'ETERS

NOTE: (_ = DIAMETER L = LENGTH --e'_,:l _kOV3.t-S +-01, 3} FIGURE 10-2. PROP-FAN STEADY PRESSURE TEST SET-UP ORIGINAL PAGE BLACK AND WHITE PHOTOGRApHr /--- COUNTERWEIGHT BLADE FORWARD TO / SIN 058 AFT VIEW #3 / #5

I

_ STEADY PRESSURE BLADE S/N OOS # 7 ,_1

I

(BLADE ANGLE MEASUREMENT SYSTEM LOCATION) FIGURE 10-3. LAP BLADE INSTALLATION STEADY PRESSURE TEST (2 BLADE} 10.2 (Continued) The Prop-Fan rotatlonal speed was measured by use of a IP pickup. The sensor was triggered by a gear mounted on the test rlg drive shaft. The rotational speed was averaged over ten revolutions.

The power absorbed by the Prop-Fan was determined by multiplylng the torque supplied to the Prop-Fan by the rotational speed. Torque supplied to the Prop-Fan was computed by accounting for the measured frlctional losses in the test rlg relative to the torque measured by the torquemeter.

Collection of blade surface steady pressure data was accomplished by utilizing a specially Instrumented SR-7L blade as i11ustrated in Figures 10-4 and 10-5, coupled wlth a Scanlvalve TM pressure measurement system.

Thirteen rows of pressure taps were located on both the face and camber sides of the steady pressure measurement blade. The pressure tap location and numbering scheme used for data acquisition and reduction are deplcted In Figures 10-6, 10-7 and I0-8. This numbering scheme differs from that In the referenced test plan. The pressure tap channels which span the blade were fabricated by bonding a thln plastic skin to a channellzed adhesive layer.

Each channel was connected to a tube embedded in the root of the blade which led out to the blade shank. The tubes were connected to the Scanivalve mounted on the dome cap of the Prop-Fan, protruding through the nose of the spinner. One Scanlvalve channel was provided for each tube. The stationary portion of the Scanlvalve contalned a pressure transducer that could be scanned by remote command to monitor one channel at a time. This arrangement allowed pressure measurements to be made at only one radial station per run.

Pressure measurements were made by masking off all the rows of pressure taps except at the section of interest. The Scanlvalve was then cycled through all channels to record the pressures at one radial station. Thirteen runs were required at each Prop-Fan operating point to obtain a complete pressure map for the blade surface at the operating point.

The Scanlvalve was enclosed in an aerodynamic falrlng to ma}ntafn a well behaved inflow to the Prop-Fan. The umbilical, which connected the Scanlvalve to the control and monitoring equipment outside the tunnel, was enclosed in a conduit with an airfoil shaped cross-section. This also minimized disturbance of the flow.

Table 10-I lists the operating conditions that were run _r_m_} the test along wlth tolerances showing the maximum variation In the parameters a11owed when testlng at different radial stations. In general, the p_=ce<Bure for setting a speclflc test condition was to set Mach number and them a{IL1ust the rotor speed and blade angle, to obtain the desired power coeff_¢_emt am<l advance ratio.

ORIGINAL PAGE BL&CK AND WHITE PHOTOGRAPH FIGURE 10-4. LAP STEADY PRESSURE BLADE {CAMBER SIDE} " AG_ OF __UA_rIp_-_ FIGURE 10-5. LAP STEADY PRESSURE BLADE (FACE SIDE) ORIGtNAL PACE BL.._CK .At,JD WHITE P l-i !.]h__ _,_ p _, ROW ! 3 )W IZ oPRESSURE PORT LOCATIONS 20' I] .NUMBERING SEQUENCE 2O OW 10 R o_w p.._ _ I ROW 8 Z0 ROW 7 Z0 ROW 6 I 2O ROW 4 ROW 3 ROW Z CAMBER SIDE FIGURE 10-6. STEADY PRESSURE BLAJ[_ • PRESSURE PORT LOCATIONS eNUMBERING SEQUENCE ROW I 3 ROW 12, ZI ROW 1 I p36 ROW 1 ROW 36 ROW 7 ROW 6 ROW 5 ROW 4 ROW 3 ROW 2 ROW !

FACE SIDE FIGURE 10-7. STEADY PRESSURE BLADE 56.0 t 142.24 52.0 132.08 48.0 "121.92 44.0 1 1.76 40.0 01.60 36.0 I, J = !, 36 16.0 '40.64 l,J- !,20 I,J= I, I '30.48 20.32 4.0 10.16 FACE SIDE CAMBER SIDE (BLADE OUTLINE MIRROR IMAGE OF FACE SIDE) 0,0' ).00 INCHES CENTIMETERS NOTE: THE VERTICAL SCALE REPRESENTS THE DISTANCE ALONG THE BLADE PITCH CHANGE AXIS FROM THE PROPFAN CENTER LINE.

FIGURE 10-8. ARRANGEMENT OF THE PRESSURE TAPS ON THE BLADE SURFACE -- PROJECTED VIEW TABLE 10-I. OPERATING CONDITIONS FOR BLADE STEADY PRESSURE TESTING (2 BLADE LAP PROPELLER) N M J /3 Cp RPM Condition Mach Advance Ratio Blade Angle Power Coefficient No. No. +J2 _+}00 ' __2 1 .01 .08 13.80' .079 1200 2 .02 .14 15.70 o .093 1200 3 .02 .15 18.780 .152 _ L20o.

4 .03 .18 21.600 .204 1200 6 .20 .88 25.650 .098 1665 6 .20 .88 30.40 ° .251 1651 7 .50 3.065 57.510 .649 1186 8 .50 3.055 54.95' .360 1190 9 .50 3.063 50.860 .108 1185 10" .60 3.066 54.980 .226 1436 11 * .70 3.055 55.00 ° .229 1685 12" .78 3.07 54.97 o .223 1840 13" .78 3.20 54.98' .112 1782 * RADIAL STATIONS 2,4 AND I0 WERE NOTRUN AT THIS CONDITION.

10.2 (Continued) After establishing the operating condition, the basic wind tunnel and Prop-Fan parameters (tunnel static temperature and pressure, Math number and rotor speed) were logged into the microcomputer, a record number was assigned for filing purposes, and the scanivalve was activated. The scanivalve then ran through a calibration sequence followed by the pressure data scan. The data were then plotted in preliminary form and reviewed. If the data contained questionable features, a second scan was performed or hand scanning of individual suspicious points was made.

Test points 1 through 4 in Table lO-I approximate static conditions of increasing power and were selected to provide information on leading edge vortex flow (there was no applied tunnel flow for these points, although there was some Prop-Fan induced flow). Points 5 and 6 were seledte¢ to - - Investigate take-off conditions: points 7 through g covered a wide range of power loading conditions at 0.50 Mach number and bracket through the design cruise power loading. Points lO through 13 were selected to Investigate transonic flow characteristics and were all at relatively low power due to rig limitations. For these points, power coefficient and advance ratio were held constant while Mach number was varied.

I0.3 Discussion and Results IO.3.l Data Reduction lO.3.l.l Data Format The pressure data were reduced to coefficient form and plotted during the test using a microcomputer system to provide the test personnel with a basis for Judging the quality of the data following each scan. This preliminary data reduction included an approximate correction (described below) for the centrifugal pumping In the tubes and channels that led from the scanivalve to the pressure taps. The pressure coefficient formula that was used is: Pc - Po Cp I 0.sp(v + Nhere: Pc - Corrected Blade Surface Pressure Po - Tunnel Static Pressure p , Air Density Vo - Tunnel Velocity Vtm Tangential Velocity at Pressure Tap Radius 10.3.1.1. (Continued) For each radial station the mid-chord radius (for a prescribed blade angle of 50.0 degrees) was used in determining the approxlmate centrifugal correction and the tangential velocity. It should be noted, however, that the radii of the pressure taps at each station vary a|ong the blade chord, and furthermore, the radius of any given tap varies with b]ade angle. In addition, the va]ues describing the chordwise distribution of the pressure taps, which were stored in the microcomputer, were nominal rather than measured values, and did not represent the precise distributions. The final data reduction at Hamilton Standard will account rigorously for these effects, and will be presented in a separate NASA Contractor Report (Reference 24).

10.3.1.2 Discussion of Problems Before revlew|ng the data, it is necessary to address three basic test problems, each of which had some effect on the data presented in this report. The problems are as follows in decreasing order of severity: l , Particle impacts resulting in failure of the tape seal over rows of dormant pressure taps or resulting in venting of individual channels.

, Crack formation In the plastic skin which seals the channels on the blade surface.

3. Reference pressure transients which interrupted the scan sequence.

Partlcle impacts on the steady pressure blade occurred on several occasions durlng the test. In most cases the impacts resulted In little or no damage, however, four cases of damage to the tape which sealed the dormant rows of pressure taps occurred and two cases of channel venting occurred due to direct partlc_e Implngement on a specific channel. The data affected by tape seal fallure were llmlted to radial stations 3, 6, 11 and 13 at Math numbers generally greater than 0.20. Channel venting due to direct particle Impact resulted on channels 26 and 29 and was repaired followlng the runs in which it occurred. The data that were compromised by Impact events will be e11minated from the final data package.

Prior to the second test run (radial station 13), cracks were found in the camber side plastic skin layer, between radlaI stations 6 and 8, and 9 and I0, on the blade. The cracks, which are believed to have been catalyzed by the appllcatlon of a cleaning solvent that "shocked" and embrittIed the plastic, were not present following the first run (radial station 5); they were found Immediately after applying the solvent. Continued blade checks showed high leak rates oo_channels 2 through 8. To correct the leak problem a series of tape strlpswas applied to the cracked areas, sealing them off and establishing acceptable leak rates. Testing continued with this conflguration and leak rates were monitored following every other run. The fix was found _.,_to_e satisfactory on all channels, although channel 2 required further repair _" as the test progressed. In general this problem is not considered to have -, compromised the data.

_- DI::_3G_AI_ PAGE rg 137 DE POOR QUAI/T¥ 10.3.1.2 (Continued) Reference pressure transients occurred during several data scans.

Fortunately, this phenomenon was quite evident in the on-line pressure coefficient p]ots, so scans that were affected were either rerun entirely or hand scanned to pick up the affected points. Therefore this problem is considered to be of minimal consequence to the test data.

10.3.2 Discussion of Result The entire spectrum of data Is summarized in Figures 10-9 through 10-21; each figure shows all of the measured pressure distributions for a specific test condition, and is in the form of the on-line data reduction. These figures are presented in order of increasing Mach number. For Mach numbers where two or more power settings were run, the figures are sub-ordered by Increasing _- power. Radial stations 2, 4, and lO are left blank for Mach numbers greater than 0.50 because no data were collected due to limited test time. Radial stations with data polnts which may be eliminated or replaced with hand logged values, for the final data package, are marked with an asterisk.

Figures 10-9 through 10-12 show the pressure distributions for the nominal static operating condition. The pressure loading is seen to increase with increasing applied power, as expected, and the presence of leading edge vortex flow Is suggested in Figures lO-ll and lO-12 by the negative pressure hump that spans along the camber side leading edge; the vortex then appears to sweep across the chord In between the 90 and 95 percent radius, resulting in very high loading in that region.

The data at these static operatlng conditions contain some inconsistencies.

For example, when running radial station 6 the data was found to diverge from the trends at the neighboring stations. The cause of this inconsistency is not certain, however, It is noted that the actual Mach number was not zero (due to the Prop-Fan induced flow) and was variable during the "static" runs (Mach No. varied from 0.02 to 0.04). Though the reason for the Mach number variation Is not known for sure, it was revealed during a daily tunnel inspection that a portion of the tunnel flow straightening honeycomb had been blown out. Subsequent running with the Prop-Fan driving the tunnel resulted in somewhat higher Mach numbers. To investigate this problem further radial station 7 was rerun for conditions 2 and 4. Comparisons of the pressure distributions or these runs are given in Figures I0-22 and 10-23. It is clear from this comparison that a repeatability problem existed for some of the static point data.

Figures 10-13 through 10-21 and Figure 10-24 show the pressure distributions for the remaining operating conditions. The data at these conditions show good repeatability. For example, Figure I0-24 shows radlal station 8 which was scanned twice at 0.20 Mach number, first on the climb up to 0.78 Mach number (in accordance with the test plan sequence), and then on the decent, approaching shutdown. The results are seen to be nearly Identical.

10.3.2 (Continued) Some of the aerodynamic effects that are apparent in Figures lO-13 through 10-21 are leading edge vortex loading at 0.20 Mach number for the take-off case (Figure I0-14), inverted leading edge pressure distributions for the low power cases at high Mach numbers (Figures 10-15, I0-18, 10-19, I0-20 and lO-21) and evidence of trailing edge shock waves at the outboard stations at high Mach numbers (evident by the trailing edge pressure jump in Figures I0-19, I0-20 and I0-21. The inverted leading edge pressure distributions, as noted above, are typical for cambered airfoil sections operating at Incidence below the design angle of attack value.

0.591 2, 0.474 RADIAL STATION 1 .

RADIAL STATION I r/R z 0.341 ¢ 1,0 00 0.2 0.4 0.6 0.8 I 0 BLADE CHORD, X/C FIGURE 10-9, LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA) .... • i i NOTES: I. TOLERANCES SHOWN DENOTE MAXIMUM 0] !__|2 PARAMETER VARIATION FROM ONE RADIAL / / S, TATION TO THE NEXT. -05. "_ Z. -'PLOTS CONTAIN DATA POINTS WHICH MAY _ I . .

BE REPLACED OR ELIMINATED. 01 ! I r I ,J' -0.5

:. • o

-z o: o _ 9 X CAMBER SIDE" ,,i _-_-_ to rACES'DE i -1.0' 0 L I ,_' -05; -,o. o_ -t,O _ 13, 0.9911 -0.1_ 5 12, 0.960 -2.0 -0]--Q t I , 0.975 0.964 -1.5 0.938 0.903 -1.0 0 0.861 -0.5 0.609 5, 0.747 0.482 -I .5 0.5_ _$, 0.591 -I .0 -:'.5 -0.5 i2.0 0_2 L_ -Z,O 0 _" -1 .S ¢_ O.S RADIAL STATION I.

r/R = 0.341 o.5 RADIAL STATION 1 1,0 0 0.2 0.4 0.4 0.0 1.0 BLADE CHORD. X/C FIGURE 10-10. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA) 0.59 1 RADIAL STATION I .

r/R -- 0.341 RADIAL STATION I 02 04 04 0,° 1,0 I_AD[ CHORD, XIC FIGURE 10-1 1. LAP SR-TL STEADY PRESSURE DISTRIBUTION (ON-LINE DATA) NOTES: FIGURE 10-12. LAP SR-7L STEADY PRESSURE DISTRIBUTION {ON-LINE DATA) l_t3 TEST CONDITION NO. 5 M N _ O.ZO = ZS.05 ° + 1.00 ° RPM - 1665 + I 0 J =, 0.88 + 0.0Z Cp : 0.098 + 0.0Z TOLERANCES SHOWN DENOTE MAXIMUM NOTES: PARAMETER VARIATION FROM ONE RADIAL STATION TO THE NEXT.

* PLOTS CONTAIN DATA POINT_J WHICH MAY BE REPLACED OR ELIMINATED, CAMBER._IDE ....

FACE SI DE 2, 0,474 RADIAL STATION I, r/R _ 0.341 FIGURE 10-1 3. LAP SR-TL STEADY PRESSURE DISTRIBUTION (ON-LINE DATA) 0. 0.2 0.4 0.6 0.11 1.0 BLADE CHORD. X/C FIGURE 10-14. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA) TEST CONDITION NO. 9 NOTES: X HAMBER SIDE J O FACE SIDE 3, 0.59 I 2, 0.474 RADIAL STATION I, r/R : 0.341 FIGURE 10-15. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA) -0.$ • T£Er CONDITION NO. 8 1 3* -I ,0_ M H = 0.50 " 54.95" _- 1.00" 0,$ RPM "IIg0±10 J = 3.055 + 0.0Z Cp = 0.360 +_ 0.0Z o _2 -0.5_ NOTES: I , TOL£RANCE;S SHOWN DENOTE MAXIMUM 0._ _ PARAM£TIER VARIATION FROM ON£ RADIAL STATION TO THE NEXT.

2. • PLOTS CONTAIN DATA POINTS WHICH MAY -0.5_ | ,I, BE R£PLACIED OR £LIMINAT£D.

o,i_ | CAMBER SIDE -o.

10 ix o.sJ O (CAMBER SIDE i FACE SI'D£ " -- - -0.5 ol .O 0.5 0.

op==,==:_-.--.--._ 6 __-- I,. o.,_ -05 °'s_ _-'*_-'-----_ _-.._ I o, o.J,.4 ,--._ _7, 0,SEt

-0, / /

o__-'.. I I __-::--.. 4*/ J__ -05 0'5= 3 -0.s] o,__ _ =, 0,474 05_ ' RADIAL STATION _, u -0.'=t " _ ___'_ r/R - 0.34I _ os___j" STATION I_ • 0 02 0.4 0,8 0,8 1,0 BLADE CHORD, X/C FIGURE 10-16. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA) TEST CONDITION NO. 7 "I'0 l MN - o.3o " t r_'T- F-T-_ - $?,51° + 1,00 ° 1 I I I I J J - _I,0S_ + 0,0Z -1.01 t _ NV/ " Cp - 0.649 + 0,02 | NOTES: I, TOLERANCES SHOWN OI"NOTEMAXIMUM _0,5t _'_,._2" PARAMETER VARIATION FROM ONE RADIAL ] J l l J J, STATION TO THE NEXT. 01 _ - _ ;, 2, * PLOTS CONTAIN DATA POINTS WHICH MAY " }

ro :::::o. i

°2.0 .0.5 FIGURE 10-! 7. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA) TEST CONDITION NO. 10 -o.5. oI_______ !* M N ', 0.60 j_ = S4,98 o +_. 1.00 ° 0.5 RPM = 14:36 + 10 -o_ 1 1 _ J = 3.066 + 0.02 Cp = 0.226 + 0.0Z NOTES: I. TOLERANCES SHOWN DENOTE MAXIMUM 0.5 PARAMETER VARIATION PROM ONE RADIAL STATION TO THE NEXT.

lO Z. " PLOTS CONTAIN DATA POINTS WHICH MAY BE REPLACED OR ELIMINATED.

FACE SIDE 0.082 -- 3.0.59 1 RADIAL STATION I, r/R = 0.341 RADIAL STATION I FIGURE 10-1 8. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA) 13" 0 _ 0.5 TEST CONDITION NO. I I M N " 0.70 - 55.00 _ + 1.00 ° RPM - 1685+ IO 1| _ J = 3.055 +_ 0.02 Cp = 0.229 +_ 0,02 o TOLERANCES SHOWN DENOTE MAXIMUM NOTES: 1.

PARAMETER VARIATION FROM ONE RADIAL STATION TO THE NEXT. 10 2. * PLOTS CONTAIN DATA POINTS WHICH MAY BE REPLACED OR ELIMINATED.

-0.5 CAMBER SIDE

o:I L x

FACE SiDE 0. S9 !

2.0.474 FIGURE 10-19. LAP SR-TL STEADY PRESSURE DISTRIBUTION {ON-LINE DATA) TEST CONDITION NO, 13 M N = 0.78 ]3 = 54.98" + 1.00 RPM : 1782 + 10 J = 3.20 + 0.02 Cp : 0.1 I Z +_ 0.02 NOTES: I. TOLERANCES SHOWN DENOTE MAXIMUM PARAMETER VARIATION FROM ONE RADIAL STATION TO THE NEXT. 0.5 2. * PLOTS CONTAIN DATA POINTS WHICH MAY lO BE REPLACED OR ELIMINATED.

CAMBER SIDE D " FACE SIDE 0,996 0.986 1 I, 0.975 0,964 0,938 0,903 0.841 6, 0.009 0.747 0.602 2, 0.474 -1.O- RADIAL STATION I, wh: -0"5_ r/R -- 0.341 O_Z o a. _ 0.5 U 1.0 0 0.2 0.4 0.6 0.8 1,0 BLADE CHORD. X/C FIGURE 10-20. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA) 4,] o i] 13" TEST CONDITION NO. 12 N - 0.78 - 54.07 ° + 1.00" -0.5, RPM : 1040 + 10 J - 3,07 _+ 0.0Z Cp = 0.223 _+.0.02 o_ I1" TOLERANCES SHOWN DENOTE MAXIMUM PARAMETER VARIATION FROM ONE RADIAL STATION TO THE NEXT, * PLOTS CONTAIN DATA POINTS WHICH MAY BE REPLACED OR ELIMINATED, -0.5 9 • 0.i 1-_

CAMBER 51D£ j

-0,5- FACE SIDE 0.$" 13, 0.986 12. 0.986 I !. 0.975 0.904 0.038 0.803 O. 801 0.809 0.747 0.082 0.591 2, 0.474 RADIAL STATION 1, r/R = 0.341 FIGURE 10-21. LAP SR-TL STEADY PRESSURE DISTRIBUTION (ON-LINE DATA) LAP (SR-7) PRESSUREDISTRIBUTION LAP (SR-7) PRESSURE DISTRIBUTION Radial St_ : 7 Record no. : 204 Radia] Sta.: 7 Record no.: 114 Beta (deg): 14.3 03-13-ig87 20:55.30 Beta (de9): 15.3 03-10-1987 2h48:01 Moch No, = 0.02 Hach No. : _02 Speed (rpm): IIQ3 Speed (rpm) : 1209 -2. 5 -2. 5 x combei side x comb_ side o-Fac_ :ide. _ io ?ace i_ide -2.0 -2.0 -I, 5 -1.5 d_ wZ z _-I,0 -_ -l.o °0. 5 _m "0. 5 kd 0.0 0.0 0.5 0.5 Io0 1.0 ........ 1 ....

o.e 1.0 0.0 0.2 0.4 0.6 o.e 1.0 0.0 0.2 0.4 0.6 BLADE CHORD. x/c BLAOE CHORD. x/c FIGURE 10-22. REPEATABILITY PROBLEM FOR STATIC LOW POWER POINT LAP (SR-7) PRESSURE OISTRIBUTION LAP (SIR-7) PRESSURE OISTRIBUTION Radial 5ta. : 7 Record no. : 205 Radial Sta.: 7 R_cord no.: [03 Beta (de9): 22.4 03-i0-xg87 18:26:06 Beta (de9): 21.4 03-13-igB7 21:04:30 NOch No. _ O. 04 Mach No. : 0.02 Speed (rpm): 120!

Speed (rpm): |203 -2. 5 -2. $ i

i i d x combe r sl e

Ix combe- side o ?ace ;ide I o Face bide -2.0 -Z.O -1.5 -].5 & z W -hn -z.o N (..I U _-O.g ua -0.5 0.0 0.0 0.5 0.5 E f 1.0 hO 1.0 0.0 0.2 0.4 0.6 0.8 0.0 0.2 0.4 0.6 0.8 1.0 BI_AOIZ C.IO_. xlc BLAOE CHORO. xlc FIGURE 10-23. REPEATABILITY PROBLEM FOR STATIC HIGH POWER POINT LAP (SR-7)PRESSURE DISTRIBUTION LAP (SR-7) PRESSURE DISTRIBUTION Rodio] 5to.: B Record no. : [B| Rodiol 5to.: B Record no.: 172 Beto (deg): 25.0 03-12-1887 21:5g: 4B Beto (de9): _6. e 03-[2m19_7 20: 2_: 50 Moch No. : 0.20 Hoch No. : 0. 20 Speed (rpm): 1701 Speed (rpm): 1670 -& 5 -Z.5 x combe side x combe side - != r_:e _Q-_ o Foce _ide -Z. 0 -Z. 0 -1.5 -65 _-].0 -l.O tJ tJ I -0.5 -0. 5 0`0 0`0 0.5 0.5 1.0 l°O 0.2 0,4 0,6 0`e hO 0`0 0,2 0.4 O.e 0`8 0`0 BLAOE CHORD. x/c 8LAOE CHORD, x/c FIGURE 10-24. ILLUSTRATION OF REPEATABILITY AT 0.20 MACH NUMBER 155/z56

II.0 BLADE SURFACE UNSTEADY PRESSURE MEASUREMENT

11.1 Test Objectives 11.1.1 Measure the unsteady pressure distribution on the surface of the SR-7L Prop-Fan blade for a range of blade angles, rotational speeds, and simulated flight Mach numbers for axial and angular inflow conditions.

11.I.2 Evaluate the effect of a wake in the propeller inflow on the unsteady pressure distributlon on the surface of a blade.

11.2 Test Procedure In preparatlon for blade surface unsteady pressure testing, the Prop-Fan was installed on the drive system as described in section 6.2, and supported as._ shown in Figures ll-l and ll-2. Table ll-I shows the conditions that were tested.

The unsteady pressure measurement blade (S/N 054) was Installed in blade position 3. For balancing purposes, the steady pressure measurement blade (SIN 009) was Installed In position number 7, as shown In Figure ll-3. As a precaution, signals from 2 strain gages located on the shank of the unsteady pressure blade were monitored throughout the test. Specla] contour matching blade stubs were installed in the remaining 6 hub arm bores. The test rlg power capabilities and tunnel time constraints necessltated conducting all testing using a 2 bladed Prop-Fan configuration, thereby permitting operation at power Ioadlngs per blade corresponding approximately to the take-off and cruise conditions of the eight blade Prop-Fan design.

The Prop-Fan was operated in the beta control mode during the entire test.

In this mode, Hamilton Standard personnel were able to change the blade pitch angle during testing by means of an Increase/decrease pitch switch located in the control room. For a fixed Mach number and a constant power supplied by the turbines, the Prop-Fan rotational speed was varied by increasing or decreasing blade pitch angle.

The blade pitch angle (B3/4) was measured and recorded for each test point.

An electrical signal proportional to blade angle was provided by means of a potentlometer mounted on the rotating side of the Prop-Fan as illustrated in Figure 5-3.

The Prop-Fan rotational speed was measured by use of a IP pickup. The sensor was triggered by a gear mounted on the test rig drive shaft. The rotational speed was averaged over ten revolutions.

The power absorbed by the Prop-Fan was determined by muItlplying the torque supplled to the Prop-Fan by the rotational speed. Torque supplied to the Prop-Fan was computed by accounting for the measured frlctlonal losses in the test rig relative to the torque measured by the torquemeter.

PRECEDING PAGE BI.NNIKNOT FILib_.O I- Z OF_ Y OR,IGINAI. PAGE BLACI'( AND ',,llIHITE P'r,KTrtY_APH FIGURE1 !-2. PROP-FAN UNSTEADY PRESSURE _ SET-UP TABLE IT-l. OPERATING CONDITIONS FOR BLADE UNSTEADY PRESSURE TEST (2 BLADE LAP PROPELLER) M J _ Cp N Condition Mach Advance Ratio Blade Angle Power Coefficient RPM No. No. +.02 +1.00 ° +.02 ±10 2 .02 .14 15.700 .093 1200 3 .02 .15 18.78 o .152 1200 4 .03 .18 21.60 ° .204 1200 5 .20 .88 25.65 o .098 1665 5A .20 .88 27.190 .15 1684 5B .20 .88 29.50 ° .20 1684 6 .20 .88 30.40 ° .251 1651 7 .50 3.065 57.51 ° .649 1186 8 .50 3.055 54.950 .360 1190 9 .50 3.063 50.860 .108 1185 10 .60 3.066 54.980 ,226 1436 11 .70 3.055 55.00 ° .229 tB85 F UNSTEADY PRESSURE BLADE FORWARD TO //S/N 054 AFT VIEW #3 / #5 STEADY PRESSURE BLADE SIN 009 _7_-"

I

{BLADE ANGLE MEASUREMENT SYSTEM LOCATION} FIGURE 11-3. LAP BLADE INSTALLATION UNSTEADY PRESSURE TEST (2 BLADE) 11.2 (Continued) Mach number was established from the ratlo of static pressure, measured four meters upstream of the Prop-Fan rotor, to stagnation pressure. Static pressure was also measured in the plane of rotation as a backup. The ratio of static to stagnation pressure was correlated wlth data taken during a pre-test calibration in order to compute the Mach number. The Prandlt-Young correction was applied to the computed Mach number to compensate for the effects of the tunnel walls and the thrust produced by the Prop-Fan.

Collection of blade surface unsteady pressure data was accomplished by utilizing a specially instrumented SR-7L blade as illustrated in Figures II-4 and 11-5, coupled with an FM multiplex data acquisition system. Twenty-six high frequency response pressure transducers were installed in two rows on the face and camber sides of the unsteady pressure blade. The p÷esgur_ --- transducer location and the numbering scheme used for data acquisition and reductlon are depicted in Figures 11-6 and 11-7.

The pressure transducers were mounted flush with the blade surface as depicted In Figure II-8. The signal and excitation wires from each transducer were connected to slgnal conditioning electron|cs located In the cuff of the blade. The slgnal wires also passed through attenuating resistors mounted on the blade root. The function of the attenuating resistors was to establish the gain for the pressure slgnals.

The unsteady pressure signals were transmitted from the rotating to the stationary field through the FM multiplex data acquisition system provided for the SR-TL Prop-Fan. The signals were monitored on a four-channel oscilloscope and recorded on a 14-track IRIG tape recorder.

The frequency response of the system was DC to 1000 Hz. Prior to the High Speed Wind Tunnel Test an evaluation program was conducted to determine the sensitivity of the transducers to temperature, strain, vibration, and centrifugal 1oadlng. The results of this test program indicated a maximum 2_ of full scale error due to temperature in the range from 0 to 130°F and a maximum .92% of full scale error due to all other factors.

Generation of the wake In the Prop-Fan inflow was accompllshed by erecting a vertlcal steel cylinder upstream of the rotor. The cyllnder was lOOmm (3.93 inches) In diameter and was located such that its centerllne intersected the Prop-Fan axis of rotation at a distance of 1.372m (54.02 inches) upstream of the rotor plane. The wake generated by the cylinder was Intended to create a twice per revolution (2P) disturbance for the Instrumented blade to pass through. Figure ll-9 shows the Prop-Fan with the cylinder _n place.

• i i i i i i i !

I I i m I C20977 FIGURE 1 1-4. LAP UNSTEADY PRESSURE BLADE(CAMBER SIDE) ORICINAL PAGE 163 8L,_C.K .A,'qD '_-_,ITE PI.K)T'b_GR,kYi, I

ORIGINAL PAGE

AND WHITE PHOTOGRAPH ~ k _ FIGURE 1 1-5. LAP UNSTEADY PRESSURE BLADE (FACE SIDE) 7, FIGURE ! 1- 6 UNSTEADY PRESSURE BLADE, FACE SIDE TRANSDUCER LOCATIONS PT14C_ T18C PT16C I

' I I

PT15C PT19C PT17C PT23C PT21C I I

I

' I ' I

PT24C PT20C PT22C PT26C FIGURE 1 1- 7 UNSTEADY PRESSURE BLADE, CAMBER SIDE TRANSDUCER LOCATIONS SHELL PLIES \ \ \ ] t / \ / \ / / \ p TRANSDUCER \ / (FLUSH MOUNTED} \ ,/ J J PETG (GLYCOL MODIFIED POLYETHYLENE TEREPHTHALATE SHEET} 0.030 THK -.j, J_ L_._ FIGURE ! 1- 8 LAP SR-7L UNSTEADY PRESSURE BLADE TRANSDUCER MOUNTING

t.V" <_,--_"_,-V;..I.?_.TY

WAKE GENERATOR FIGURE 1 1-9. UNSTEADY PRESSURE TEST SET-UP WITH WAKE GENERATOR ORIGIN'A/ PA_[' 6LACK AND _'HITE PHO'IOGRAPId = 11.2 (Continued) The complete range of Prop-Fan operating conditions that were tested is given in Table ll-I. Because of time limitations, the range of conditions tested was reduced for the zero degree inflow cases with and without the cylinder.

In addlt_on, transducer wire problems resulted in intermittent and nonexistent signals on several sensors. The resulting series of test condltions for which signals were recorded is given for each transducer in Table ll-II.

11.3 Discussion and Results To provide an illustration of the data collected during the test, examples are given in Figure ll-lO of the periodic variations In pressure (with corresponding frequency spectra) which were measured on the camber side ok - the blade at the 90% radius, 56% chord point (pressure transducer number PTI6C). Here the three inflow cases are compared for the Prop-Fan operating condltion defined below.

Mach number, MN - 0.20 Advance Ratio, 3 = 0.883 Power Coefficient, Cp = 0.250 Blade Angle, 8 = 32 ° This operating point Is representative of the Prop-Fan take off condition.

The pressure versus time plots at the left in Figure II-I0 were obtained from a sIgnal enhancing waveform analyzer. Sampling was Inltiated by the recorded once per revolution plp signal and waveforms from I024 revolutions were averaged. Thus the repetitive portion of the pressure waveform is enhanced and the random part Is suppressed. The spectra, shown at the right In Figure ll-lO, were obtained vla digital Fourier transform analysis.

Successive time slices were transformed and averaged for 4.8 seconds. Each spectrum conta|ns 400 frequency points spaced linearly from 0 to 500 Hertz.

Prel|minary Interpretation of Figure ll-lO is as follows: In interpreting the waveforms, a transducer can be considered to be scanning the inlet flow as it rotates. Since the blade position Is known as a function of time, the time axls can be converted to angular position. The six and twelve o'clock positions are indicated on the bottom trace. For the trace at the top of Figure ll-lO, representing clean inflow, the signal level should be low, corresponding to a low distortion level. However, a small slnusoldal component can be seen In the waveform and spectrum that must be caused by a residual flow angularity In the tunnel. This can be consldered a background level and must be subtracted from the data for the 3° angular inflow and for the cylinder wakes.

F-.

<7

:+

O_ _i u +, ...i Z

:+,-+

+<

_zl-

o+

1,11:3 .-12:; I+,

I<_

I++

li: l-- h o !- a _ I,-,, m-_ z <_ 0 < "I- m" o < (,,3 W I-_ I"

+i l_-"tl._ . _,_ _ a

hi li,i ,.,.I < lli ">a < ": _i I-- - _ +_ +7<©C+p++ ++++r+,+, +- N N N _, .n. _ r_l I_ _-- r_ ,_':" / U I-. >" I- u I-_ _" J w Z __ till. _. ;_ a. J "_. _Ja ............... _t_,r ...................... [ b.

W '- zo _ I .--. _. [---.-.-.-.-.-.-.-.-_-- .J -ou.... _ ............... ._. _ ................. "' "._.--"_--_ ...... ! ............

re (Jl ........ :-.'.:: ........ _ o .......... ................ o < i_. ................... _'-_.. _ i ,.,Z - ,. - _ _°'"_ n, %. . o o _.

0 ISd _l_l'_GS3_d _" "o _ - >.

<

o o i ¢

I-- Z C_ < u u E P_ _ "I" !- >" o >. ..... _I_ .z" _ Jou ° .........

_ < uJ"_ ..s- _/_.

.J zo z0 o > zJo_ L. --_. _/_ _" _" < i'l' OI _ Olu < X .i ...... ......... -I _r -- - Ld T _ _ _ ® o o o, . o/ _. /,., ; m ! ..... °o _ I_ ,' o _ _- _o_'_ o o n U _ _ _ _; _ <_o + , + ' + . hi !-- N U) " Z I" '_ < ISd "lVNIINON INOM.-I NOI.ILVlElVA 31::ll'tSS3_d eIOV'1fl 11.3 (Continued) In the data for the 3° angular inflow, the angle of attack seen by the instrumented blade should be nearly a pure sine wave at the once per revolution (IP) frequency. Simplistic analys_s wou}d _ndlcate that the blade pressure response should also be sinusoidal. The waveform and spectrum show that this is far from true. Figure 11-11 illustrates the terms "advancing" and "retreating" for angular inflow. Further evaluation of the data at other positions on the blade is required to identify the source of this non-slnusoidal behavior.

For the data with the cylinder wake, the blade pressure should respond with a pulse each time the blade passes through a wake at the top and bottom of the revolution. This behavior Is observed in the bottom trace, but the pulse magnitudes are surprlsingly different at the top and bottom positions_ ....

Another Interesting feature of the data for cylinder wakes is the oscillating response after the wake pulse.

Slnusoldal response was observed on the pressure (face) side of the blade In all cases examined for angular inflow conditions.

Sinusoldal response was also observed on the suction (camber) side of the blade under low 1oadlng conditions. However, under high loading conditions, non-slnusoldal behavior is present. The non-slnusoldal response appears to be a result of leading edge and tip vortices which may be distorting the response. Another possibility Is the formation and breakdown of the vortices as the angular Inflow or wake inflow modulates the angle of attack.

An analysis of the blade unsteady surface pressure data will be presented in a separate NASA Contractor Report (Reference 25).

Z RETREATING BLADE INFLOW i, IL % _ w ADVANCING BLADE FIGURE 11-1f. ILLUSTRATION OF THE TERMS "ADVANCING" AND "RETREATING" FOR ANGULAR INFLOW 173/174 12.0 CONCLUSIONS The High Speed Mind Tunnel Test has provided an extensive evaluatlon of the operating characteristics of the SR-7L Large-Scale Advanced Prop-Fan. All the test objectives, set forth in the Plan of Test 267X-135 Rev. D, regarding acquisition of data were accomplished and I02 hours of operating experience were attained. No problems were uncovered that would have been considered an impedance to the planned follow-on PTA Flight Test. ONERA drive system constraints and testing problems precluded running all desired test points, however, those polnts successfully run did provide a wealth of aerodynamic performance, structural dynamic, steady and unsteady blade pressure Informatlon. Further areas of investigatlon are indicated that should ultimately result in highly accurate aerodynamic design and performance prediction methods. The conclusions and recommendations derlved from each phase of the High Speed Wind Tunnel Test are presented in the follow_ng -- - sections. Two separate low number NASA Contractor Reports will be published providing a more detailed evaluation of the blade surface steady and unsteady pressure tests.

12.1 Blade Structural Dynamic Evaluation The SR-TL Prop-Fan was found to be free of high speed blade flutter over the entire operatlng envelope tested. All the measured blade surface and blade shank strains were below the a]lowables set prior to testing. These allowable levels were set to avoid accumulation of fatigue damage to the blades, therefore, no fatlgue damage to the blades was incurred.

Reasonable correlation was found between the measured and analytically predicted IP blade bending strains for the SR-7L blade. Results confirmed that IP strain peaks inboard on the blade and lessens near the tip. Also, blade strains were found to Increase with power and Mach number (all other variables held constant).

The Interpretation of IRP (infrequently repeating peak) strain values was made difficult due to a significant amount of high frequency (>25P) noise in the signaTs. In generaI, this type noise Is easfTy filtered out, however, this could not be done for the zero inflow angle data due to the low signal response levels. The angular inflow (3 °) strains were determined from spectral analysis, the amplitudes of which were not affected by "noise" outside the frequency range of interest.

Signiflcant 2P blade vlbratory response was also measured for the 3° angular Inflow case. The amplitude of the 2P response averaged approximately 25% of the IP component response. It is concluded that at least a portion of the excitation force driving the 2P response was generated by the test rig drive system. Twice per revolution vibration is characteristic of a shaft with a universal Joint. It Is also noted that nonlinearities in the aerodynamic excitation, possibly due to observed vortex loading phenomena or tunnel inflow irregularities, are causes of higher order excitation.

12.1 (Continued) Because RPM traces were not available for the data reduction effort, the harmonic content of the blade strains was determined from a "non-speed corrected" spectra] analysis. The spectrum values were compared to values obtained from calculations of amplitude of the IP bandpass filtered signals.

It was concluded that the blade strain magnitudes from the spectral analysis were withln I0% (low) of the actual magnitudes.

For comparison purposes, IP strain predictions were also made for the angular inflow cases. Due to test constraints, angular inflow test data collection was limited to a 2 bladed Prop-Fan configuration at 3 ° inflow angle.

Predictions were 16 to 31% higher than test data for the inboard response (flatwlse shank moment and radial bending) and up to 13% lower than test for the outboard bending response. Th|s is consldered reasonable Correl_tton, especially when it is considered that the measured values were perhaps up to 10% lower than correct. (See prevlous conclusion). As w|th earlier testing of the SR-7A 2 ft. diameter aeroelastlc model, correlation of root response was marginal for the trailing edge, shear and chordwlse gages (Reference 26).

12.2 Aerodynamic Performance Evaluation Measured aerodynamic performance of the SR-7L Prop-Fan corresponded well with analytical predictions for the four blade configuration over the entire range of points tested. Similarly, good agreement between measured and predicted performance was found for the eight blade configuration at Mach numbers of .70 and .73. However, performance of the elght blade configuration was slightly underpredlcted at .50 Mach number. The characteristic shape of the LAP performance curves were slmiiar to those observed for the SR-7A aeroelastic mode] wlnd tunnel tests (Reference 27).

12.3 Blade Surface Steady Pressure Measurement Steady pressure distributions were successfully measured on the blade surface at a11 of the 13 radial stations for Mach numbers of 0.03, 0.20 and 0.50, and at all radial stations except 2, 4 and lO for Mach numbers of 0.60, 0.70 and 0.78. Subsequently, an uncertainty analysis was performed, which demonstrated that the measurement errors and uncertainties involved In th|s test, were acceptable.

Studles of the sensitivity of the correction for centrifugal 1oadlng on the column of air In the blade's pressure tap channels showed that the assumptions used In processing the data were valld.

During operation at approximate statlc rotor conditions, the Inflow Mach number was 0.03 +0.015, where +0.015 Is the maximum statlon-to-station variation In the-Math number. -The pressure distributions were found to be very sensitlve to Mach number for these conditions, so some respective station-to-station pressure dlstributlon inconsistencies exlst In these data.

12.3 (Contlnued)

The chordwlse pressure loading Is found to moveaft with increaslng relatlve

Mach number for the 0.60, 0.70 and 0.78 Mach number cases.

Evidence of tlp edge and leadlng edge vortex flows were found in the statlc rotor data and the 0.20 Mach number data.

The inverted leading edge pressure distributions observed during the low power high Math number conditions, are typical for cambered airfoil sections operating at Incidence below the design angle of attack value.

Evidence of trailing edge shock waves Is present at the outboard radial stations In the 0.70 and 0.78 Mach number data.

12.4 Blade Surface Unsteady Pressure Measurement Unsteady blade surface pressure data were successfully measured over the followlng range of conditions" Angular Inflow (3 °) 0.02 ! MN _ 0.70 Unlform Inflow (0 °) 0.02 < MN < 0.50 Inflow with Wake 0.03 < MN < 0.50 The uniform Inflow data shows evidence of distortlon. This appears to be a result of test section Inlet asymmetry.

A The angular Inflow and wake data clearly shows unsteady pressure response.

dominant once-per-revolutlon response is evident in the angular Inflow data while the wake data i11ustrates twlce-per-revolution response as the Instrumented blade passes the wake generating post.

Slnusoldal response was observed on the pressure (face) side of the blade in all cases examined for angular Inflow conditions.

S1nusoldal response was observed on the suction (camber) side of the blade under low loading conditions. However, under high loading conditions, non-slnusoldal behavior Is present. The non-slnusoldal response appears to be a result of leadlng edge and tip vortices which may be distorting the response. Another posslbillty Is the formation and breakdown of the vortlces as the angular Inflow or wake Inflow modulates the angle of attack.

177/178 LIST OF SYMBOLS _mm,-=__# r']_ ..... ,, LIST OF SYMBOLS A area Incremental forward centerbody area at tap N AN AM Incremental aft centerbody area at tap M AF blade activity factor = 6250 f"°(b/D) X3dX Hub/tip & area weighting factor BHP brake horsepower, KW BF buoyancy force, N DCBT CCBDT centerbody drag coefficient without blades = qo (Ac,) DcB CCBDW centerbody drag coefficlent with blades qo (AcB) Ds CSD sp|nner drag coefficient = qo (As) Cp power coeffic|ent = poN3D s thrust coeffIcient - CT poN2D" TApp - BF net thrust coefficlent = CTNET poNZD 4 C speed of sound, m/sec D drag, N D dlameter, m EF excltatlon factor = 9(VT/644.8)Z(p/po) F force, N IRP Infrequently repeating peak V advance ratio = 60 -- ND M Mach number N rotational speed, RPM P power, watt P pressure, N/cm 2 PA pressure forces in the form (P-Po) Area, N Po freestream statlc pressure, N/cm 2 PN, PM static pressure at tap N, M PT total pressure, N/cm z PSTAG stagnation pressure, N/cm 2 Q torque, N'm dynamic pressure, N/cm z qo r radius, m r/R fractional radlus SHP shaft horsepower T thrust, N TSTAG stagnation temperature, °K Ts static temperature, °K T_ total temperature, °K V velocity, mlsec blade angle, deg q effIclency mass density, Kg/m 3 inflow angle, deg thrust p(d_sc area)V 2 dlsc area w tunnel cross-section area Subscrlpts APP apparent B balance BP back pressure CB centerbody CBT centerbody tare corrected C, COR - FL funct|onal losses MEAS measured N number NET net free stream S spinner t tangential T true TH thermal effects TNET net thrust lsl/ls2 REFERENCES mmmJ,_m ....fl_'_J REFERENCES I. J.B. Whitlow and G.K. Sievers, "Fuel Savings Potential of the NASA Advanced Turboprop Program," NASA TM-83736, 1984.

2. J.F. Dugan, Jr., NASA LeRC, B.S. Gatzen and N.M. Adamson, United Technologies, "Prop-Fan Propulsion - Its Status and Potential", SAE Paper 780995, November 1978.

3. "Cost/Benefit Tradeoffs for Reducing the Energy Consumption of the Commercial Air Transportation System, Douglas Aircraft Co; NASA Ames Contract NAS2-8618, NASA CR-137925 Summary Report, June 1976.

4. "Study of the Cost/Benefit Tradeoffs for Reducing the Energy Consumption of the Commercial Air Transportation System", Lockheed-CalifOrni_ - - - Company, NASA Ames Contract NAS2-8612, NASA CR-137927 Summary Report, August 1976.

5. "Energy Consumption Characteristics of Transports Using the Prop-Fan Concept", Boeing Commercial Airplane Company, NASA Ames Contract NAS2-9104, NASA CR-137938 Summary Report, October 1976, 6. "Study of Unconvent|onal Aircraft Engines Designed for Low Energy Consumption", Pratt-Whitney Aircraft, NASA Lewis Contract NAS3-19465, NASA CR-135065 Final Report, June 1976.

7. "Study of Cost/Benefit Tradeoffs for Reducing the Energy Consumption of the Commercial Air Transportation System", United Airlines, NASA Ames Contract NAS2-8625, NASA CR-137891, June 1976.

8. "Study of Unconventional Aircraft Engines Designed for Low Energy Consumptlon", General Electric Co., NASA CR-135136, December 1976.

9. "Fuel Conservation Merits of Advanced Turboprop Transport Aircraft", Lockheed-Californ|a, NAS2-8612, NASA CR-152096 August 1977.

lO. NASA Tech. Mem. 83736 "Fuel Savlng Potential of NASA Advanced TurboProp Program", September 1984.

11. D.C. Mlkkelson, G.A. Mitchell, and L.J. Bober, "Summary of Recent NASA Propeller Research", NASA TM-83733, 1984.

12. NASA CR 3505 "Evaluation of Wind Tunnel Performance Testing of an Advanced 45 ° Swept Eight Bladed Propeller at Mach numbers from .45 to .85", Hamilton Standard Dlvlslon, Windsor Locks, CT, March 1982.

13.

C.L. DeGeorge, J.E. Turnberg, H.S. Wainauski, "Large-Scale Advanced Prop-Fan (LAP) Statlc Rotor Test Report", Hamilton Standard Division UTC, Windsor Locks, CT (NASA CR-180848).

14.

W.E. Su111van, J.E. Turnberg, J.A. Vlolette, "Large-Scale Advanced Prop-Fan (LAP) Blade Design," Hamilton Standard Division, _indsor Locks, CT, (NASA CR-174790).

15.

R.A. Schwartz, P. Carvalho, M.O. Cutler, "Large-Scale Advanced Prop-Fan (LAP) Pitch Change Actuator and Control Design Report", Hamilton Standard Division, Windsor Locks, CT, January 1986 (NASA CR-174788).

16.

M. Soule, "Large-Scale Advanced Prop-Fan (LAP) Hub/Blade Retention Design Report," Hamilton Standard Division, Windsor Locks, CT _ _ (NASA CR-17¢786).

17.

B. Huth, "System Design and Integration of the Large-Scale Advanced Prop-Fan, "Hamilton Standard Division, Windsor Locks, CT, August 1984, (NASA CR-174789).

18.

Pierre, M. "Characterlstlques et Possibillte de la Grand Soufflerie Sonlque de Modane-Avrleux," ONERA Technical Note 134, 1968.

19. Masson, A. "Essair D'Hellces dans ia Grande Soufflerle de Modane-Avrieux," ONERA Technical Note 161, 1970.

20.

Pope, A., "Wind-Tunnel Testing," John Wiley and Sons, 1954.

21.

R.M. Reynolds, R.I. Sammonds and G.C. Kenyon, "An Investigation of a Four Blade Single-Rotatlon Propeller in Combination with an NACA 2-Serles, D-Type Cowling at Mach numbers up to 0.83", NASA RM A53B06, April 13, 1953.

22.

H.G1auert, "Airplane Propellers, Body and Wing Interference", Volume IV, Division 2, Chapter VIII, Aerodynamic Theory, W.F. Durand, editor Juluis Springer (Berlin), 1935 (Dover reprint 1963).

23.

"BESTRAN User's Guide, Auxiliary Program ST570", D.C. 3ennings, December 14, 1977.

24.

Bushnell, P.R., "Measurement of the Steady Surface Pressure Distribution on a Single Rotation Large Scale Advanced Prop-Fan Blade at Mach numbers from 0.03 to 0.78", (NASA CR-182124).

25.

Bushnell, P.R., Gruber, M.E., and Parzych, D.J., "Measurement of the Unsteady Surface Pressure Distribution on a S|ngle Rotation Large Scale Advanced Prop-Fan Blade at Mach numbers from 0.03 to 0.70", (NASA CR-I82123).

26.

D. Nagle, S. Auyeung, 3. Turnberg, "SR-7A AeroelastIc Model Design Report", Hamilton Standard Divls|on UTC, Windsor Locks, CT (NASA CR-174791).

27.

G. Stefko, G. Rose, G. Podboy, "Wind Tunnel Performance Results of an Aeroelastlcally Scaled 2/9 Model of the PTA Flight Test Prop-Fan", AIAA-87-1893.

lss/ls6 ,i

APPENDIX A CHRONOLOGICAL HISTORY OF TEST

APPENDIX A

CHRONOLOGICAL HISTORY OF TEST

Testing of the LAP Prop-Fan started on February 20, 1986 and continued until

April 9, 1986. The test was terminated at this point following the discovery

of severe fretting corrosion in the ONERA drive shaft retent|on area. During

this time period, 55 hours and 3 minutes of test tlme were accumulated.

Testing resumed on February 27, 1987 and contlnued until March 19, 1987. The

intent of the second phase of testing was to collect blade steady and

unsteady surface pressure data utlIizlng a 2 blade configuration. Testing

was successfully completed after accumulating an additional 47 hours and lO

mlnutes of test time.

The following tabulation provides a chronological history of the entire High

Speed Nind Tunnel Test conducted at the ONERA facil|ty In Modane, France.

I-- I-- ,- I I I I t I I f I I I I I I I t_ 4_ I-- ¢l .-I I- I 4-1 _0- I ¢J _,- O.

i--I "0

¢

I "0 4_ O_ l- ¢J _ _ e'_ Oi i (J I/1 _ 0 _ e'- e'_ e-- I- o,- _ 0 "1 ¢1 • "0 "0 U- n_ I .,- e_ ,"14=, Ill klJ 0",- _ e-- Z Z q_ ko oo oo CO CO O0 O0 CO

& ,_" & ,_ J

e_ ed ex; o4 eq e_ eq __ L e- m3 r-j _'3 _'3 .,-.

4-) °p.- °.

I i I I I {.)

I I I I I I I I I --I _=.

o

v I G# S- _ Z _:I ¢.- I--" _ o 4.-; i."-

l.l.. o

¢-,j

Oi tel

I _'-

>-

i'm .,- I _:I e- W

o "I_ U

Z d_ _ _- • 0 4.-) z 4.o _ __ ._-- ° , ° "0 _.n 4...)

•-I _13 aJ llJ I_ 0 It I •i-J .._ I-. _,1 c

0 r- 4-., _

---

.I,-._ "I_ 0 _ 4-_ 0 "_- r- _- u 0 _-

0 0"o _ U

•"o _ ,- 0 0 0.-_

_ 0 •,-- LJ L.I c13 (I_ > Q.

0 0 _' -1-,, _,_ ._ 0 •_ ,--- -I-_ e"l, _ 4.o _.. ,.-- ,,,I--I _ •

0 m

e.- e- ,I-_ _'m e" _I_ _-- II l..- _ _ .-m z e- i-- _.) r- e- e- e- ,,.-

,4 _0

(M ('_ 0"1 I- e_ e- e- e- i, o1-- {.- .p.- 00 C_ ¢M L#I ,c/" _I- O'_ L/1 =E Lr) O .o .o p_

.:: ;4

E

u v (13 r- o.

I I I in 4.m "0 O rt_ (J _[j ,-- v _,1 ..E Z u': ,--- }l "O O -T- rt_ I--

P- a-_ IJ 0 0

(-_ _0_ O0 O _1. (D O • GI -1-) _ .4_ _-O _-. _L) U • ,- I II O :_ (-" _S-O aO 0 0 _.. $- O _- z ._,- :3 I r- (]j .,.-- _.

I _IJ II ._- I _" "0 ::3 O C: i: _ z O O _--'_ 0 C ..l- "_XD _" "_ "t_ rtS 0-_

,., O g:c

_J :_ .._ -_ _- 0 e- > O .E'I_ O P_" _" -I-_ ._ _. f,.. {= i--i

o = "t_ u

I U'IO (1; - (fl P- N E U .fa

>

E'Q "_ _-- _1_0 > ,-.- E: _=,-- _-- I oP :3 t--- _ O rl_J_ _:_ 4-) C_.(_ _0 _ %.

_- ..-- _y .4-- • _- 01.--" C_ _E ,J,-: c c i_ _- 0 _._ ,_ ._, (f) L_J _" _" J:l 0'1 E .Z 3 :_ 3 • z O :Z:O CO 00 00 0(3 0{3 L_ .4= (-,,J

S

::3 "--.,I= I- U _=_ $,., _.. I.- _- $.... $- t._ I- i- 1.4_ _ _ 1.4. 1.4..

A

E

Lo o'I 0 0 eJ 0 °° °° _ .° CO CO CO ¢v I ¢v I ¢v_ F-- r-- F-- _'_ v m_- c_ o I I

I °" I "° _i Ji ""

o

c- O t_

_ _ _o

tlt2 4.._ ,_- 0 0-

>o

tn XD °_ CO u_ "oZ 0

:_ ' % _ o = ®_-

q-- -i=_v c- t_ O °_

_ _: g _= •

-0 .0 c- c- _j c- tn c- c- ,-- 0 tO cO CO O0 Cn 0"_ 0"_ e- e- e- e- e-- e- e- e-- e-- e-- t2 I.- $.. l-- X X .r- _- u_ ,o t, v u _0 oO O_ r-_ c_ <

6 _i

0,-.

I'-- • e- 4-1 ._ co o'D o"1 o'1 CO c_ 0 0 0 0 0 0 0 0 0 oo oo .o .o .o oo L,O _ 1.0 UO _ 1.0 kO I_.'1 ul LO 1.0 e- .v,.- _.Jv I-- =D I.n O- ! I I t I I r_ ,e- t#1 Q_ > ,-- 0 .0 E • -_ '" "_ 0 0 _ "_ _ 0 I _- O0

= _ o-_

e,,.'

I-- 0 _ O' _ _ 0 0 _- c_ c'- .:I .,- G) E =1 r- 0 > 0 _ _1_ r_ 4-1, l,-- v O0 _0 O0 I-- o..I ,c..I

" a_ '--

"_ "Z "Z 'Z

0,. 0.. 0.. 0..

r--j _ i--j, < < < < < ,,_ < P I--- .r-- .<","1 o_ t_ 0 0 0 0 0 0 0 0 0 o, ,o ..

.o ,o ,, ._ ,, .o .. o, I.n ut_ Lt_ v i I I I I i i =3 t-- e..

0 _n _L_ v .4-1.1-_ r_ ."s 4_ 0 > I'- t_ _0 t:e" ::3 _ • 0 Oi e- _..

1-- ® 0 _ O_ 4-_'_ _ _.

e- 4_ Q_ l (J U _ I e-- _1_ #J "_ I 4._ e- I c- O 0 "_ 4_ I 4-} In e- _. e- oo 0"_ l 4_ • 4.;, tn I1 e- _ o 0 0 z I _ >, ,.-._ 0 ClJ 4.J ,'_ O_- U o_=. ,_-.

e- 0 _ > "aO 0"_ "_ _l_ oJ • "1 H _ 4..z"l_ ..i..z r_ _. e.. _l_ _j_ 0 _.-_ e.- ;_ z 4_ r- e- I e-- o'_ e" e" 4_ r_ cL_ Z e- oe- I_ e-e-

= o o

04-' r_ p_ aO aO a0 C_ I,---- r,-, C;

' c4

N $..

%.

f,- f,.. I,.. L l- I,- $.- L f,-- $-- e'J CI_ QJ CI_ I.i-- h I, 1.1_ I._ I,_ ,j °, u o'_ o') o) 0 r_ cO _ U =) 0 ,, ,, ,. °.

0, c- O L_ v I I -- r- (_ > > _ _..,- ,__ _ c" _=0 _,-- _" C_ t_ > _ _- _-_ 0 "0 =o_ 4-) _ _=o 0 i.-i -.s >

_o _-- ,_

,-- _. _ _:I _- ._ -_,.-- c 0.._ _ _cO

o_ ">'

u'_ e" GI 4-_

-_,_ O0 _

r_

"-g8 =

r._ 0 c O0 ,_0 u>_u _0 _- _.

_q>- ,.- ,o 0 0

__E '-

_v

¢_j t.- I I "1o I-- l-.-- c U I,-- e- I'- I Q_ • 4-P ¢'_ Q.

_.j _ _ 0 °f,,- !I_ e-- e- _ E oo co r,j _.j f_ 1, ReDort No. 2. Government AcceSsion No. 3. Recipient's Catalog No.

NASA CR 182125

5. Report Date 4. Title and Subtitle

JULY 1988

LARGE-SCALE ADVANCED PROP-FAN (LAP)

6. Performing Organization Code

IIIGH SPEED WIND TUNNEL TEST

73030

7. Author(s) 8.

Performing Or_nizauon Report No.

William A. Campbell Peter Arseneaux

HSER 11894

Harold S. l_ainauski

10. Work Unit No.

9.

P_f_ming Orgini_tion Name end Address ._35-03-01

HAMILITON STANDARD DIVISION

11. Contract or Grant NO.

UNITED TECHNOLOGIES CORPORATION

ONE HAMILTON ROAD

NAS3-23051

WINDSOR LOCKS, CT 06096

13. Type of Report and Period Covered 12.

Sponsoring Agency Name and Addrm

CONTRACTORREPORT

NASA Lewis Research Center

14. Sponsoring Agency Code 21000 Brookpark Rd.

Cleveland, Ohio 44135

15.

SuD_emen_ Note=

Project Manager: J. Notardonato, Advanced Turboprop Project Office

NASA Lewis Research Center

Cleveland, OH 44135

16.

Abs_a_

High Speed Wind Tunnel testing of the SR-7L Large Scale Advanced Prop-Fan (LAP) is herein

reported. The LAP is a 2.74 meter (9.0 FT) diameter, 8-bladed tractor type rated for 4475 KW

(6,000 SHP) at 1698 RPM. It was designed and built by Hamilton Standard under contract to

the NASA Lewis Research Center. The LAP emp_ys thin swept blades to provide efficient propulsior at flight speeds up to Mach .85.

Testing was conducted in the ONERA SI_ Atmospheric Wind Tunnel in Modane, France. The test

objectives were to confirm the LAP is free from high speed classical flutter, determine the structural and aerodynamic response to angular inflow, measure blade surface pressures (static and dynamic) and evaluate the aerodynamic performance at various blade angles, rotational speeds and Mach numbers.

The measured structural and aerodynamic performance of the LAP correlated well with analytical predictions thereby providing confidence in the computer prediction codes used for design.

There were no signs of classical flutter throughout all phases of the test up to and including

the 0.84 maximum Mach number achieved. Steady and unsteady blade surface pressures were

successfully measured for a wide range of Mach numbers, inflow angles, rotational speeds and blade angles. No barriers were discovered that would prevent proceeding with the PTA (Prop-Fan Test Assessment) Flight Test Program scheduled for early 1987.

17. Key Woads(Suggested by Author(s))

Advanced Turboprop T_is docu_ wri11 remain under distributi(

Prop-Fan

limitation uat_I July 1989.

Energy Efficient Propeller

High Speed Wind Tunnel

lg. Security Oauif. (of this report) 20. Security Qa_if, (of this I_l t Z_t. No. of Pages 22. _ice"

UNCLASSIFIED UNCLASSIFIED [ 193

_ NASA-C-16@ (Ray. 10-75)