Under Distribution Limitation

Until December 31, 1987

NASA CR-175025

NASA

ANALYSIS AND TEST EVALUATION OF THE DYNAMIC

STABILITY OF THREE ADVANCED TURBOPROP MODELS

AT ZERO FORWARD SPEED

By

Arthur F. Smith

HAMILTON STANDARD DIVISION

UNITED TECHNOLOGIES CORPORATION

WINDSOR LOCKS, CONNECTICUT 06096

December, 1985

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135

Contract NAS3-22755

I, Report No. 2. Government Accmion No. 3, ReciDient's Catalog No.

NASA CR- 175025

4, 5. Regort Date

December, 1985

Title and Subtitle

Analysis and Test Evaluation of the Dynamic

Stabilityof Three Advanced Turboprop

Models at Zero Forward Speed

Author(s)

Arthur F. Smith*

Performing Organization Name and Address

Hamilton Standard Division

United Technologies Corporation

Windsor Locks, Conn. 06096

12. Sponsoring Agency Name Ind Address

National Aeronautics & Space Administration

Washington, D.C. 20546

6. Performing Organizltion Code

8. Performing Organization ReDort No

HSER 11054

13.

10. Work Unit No.

11. Contract or Grant No

NAS 3-22755

Type of Report and Peciod Covered

Contractor Report

14. S¢ontoring Agency Code

15. _Di)lementary Notes

Final Report, Project Technical MonRor, O. Mehmed, NASA Lewis Research Center, Cleveland, Ohio 44135

*Now with: Hirock Corporation, i Main Road, Granville, MA 01034

16. Abstract

Results of static stability wind tunnel tests of three 62.2 cm (24.5 in) diameter models of the Prop-Fan are

presented. Measurements of blade stresses were made with the Prop-Fans mounted on an isolated naceDe

in an open 5.5 meter (18 foot) wind tunncl test section with no tunnel flow. The tests were conducted in the

United Technology Research Center Large Subsonic Wind Tunnel. Stall flutter was determined by regions of

high stress, which were compared with predictions of boundaries of zero total viscous damping. The

structural analysis used beam methods for the model with straight blades and finite elemet_ methudu for the

models with swept blades.

Increasing blade sweep tends to suppress stallflutter. Comparlsions with similar test data acquired at

NASA/Lewis are good. Correlations between measured and predicted criticalspeeds for allthe models are

good. The trend of increased stabilitywith increased blade sweep is well predicted. Calculated flutter

boundaries generally coincide with tested boundaries. St.allflutteris predicted to occur in the third (torsion)

mode. The straight blade test shows third mode response, while the swept blades respond in other modes.

J

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

Advanced Turboprop Propellers

Prop-F an

Energy Efficient

Blade Stall

Stall Flutter Tests

19. Security Classif. (of this reDortl

Unclassified

18. OistributiOn Statement

Unitl December 31, 1987

NASA-C-168 (Rev lO-TS)

FORENORD

All of the testing reported herein was performed at the United Technologles

Research Center facilitles in East Hartford, Connecticut, under the dlrectlon

of personnel from Hamilton Standard, a Division of United Technologles

Corporation. The assistance of UTRC personnel in the performance of thls

testing Is gratefully acknowledged. This work was accomplished under contract NAS3-22755 for the NASA-Lewis Research Center In Cleveland, Ohio.

Mr. Oral Mehmed of the NASA-Lewls Research Center was the Project Technlcal

Monitor for this contract.

The data reduction was performed by Mr. Donald J. Marshall, and the analysis

and reportlng was conducted by Mr. Arthur F. Smith. Mr. Bennett M. Brooks

was the Hamllton Standard Project Manager.

FRECEDING PAGE BLA_"_}{ NOT FILML{D

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TABLE OF CONTENTS

ABSTRACT ...........................................

FORENORD ...........................................

TABLE OF CONTENTS ..................................

SUMMARY ............................................

SYMBOLS ............................................

INTRODUCTION ................. "......................

DESCRIPTION OF EXPERIMENTAL PROGRAM ................

2.1 Model Description .............................

2.2 Test Models ...................................

2.3 N1nd Tunnel Facility ..........................

2.4 Propeller Dynamometer .........................

2.5 Model Instrumentation .........................

2.6 Test Procedures ...............................

2.7 Operating Conditions ..........................

2.8 Data Reduction ................................

ANALYTICAL TECHNIQUES ..............................

3.1 Method Description ............................

3.2 Calculated Instablllties ......................

TEST DATA EVALUATION AND COMPARISON NITH CALCULATIONS

4.1 Response Frequencies ..........................

4.2 Total Stress Results ..........................

4.3 Spectral Analyses .............................

4.4 Modal Response for the SR-3 ...................

4.5 Comparison with NASA-Lewls Tests ..............

CONCLUSIONS ........................................

RECOMMENDATIONS ....................................

REFERENCES .........................................

TABLES .............................................

FIGURES ............................................

APPENDIX A - TOTAL VIBRATORY STRESS PLOTS .... _.....

APPENDIX B - SR-3 STRESS PEAK FREQUENCY TABULATION...

APPENDIX C - SR-3 CAMPBELL DIAGRAMS ................

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SUMMARY

Static stall flutter tests were conducted in an unattached open 5.5 meter

(18 foot) test section of the UTRC wind tunnel on three Prop-Fan models.

These models are designated the SR-2, SR-3 and SR-5 with the blades

characterized by increasing sweep, from the unswept (straight) SR-2 blade to

the highly swept SR-5 blade. The tests were conducted at zero flight speed,

over a large range of blade angles and rotational speeds (RPM), Including all

areas of deep stall. Blade vibratory stress measurements were recorded for

all operating conditions. Extensive analysis of these data was performed.

Perhaps the most significant test result seen is that increased blade sweep

is beneficla] In suppressing the hlgh stress which Is indicative of sial]

flutter. The unswept SR-2 model is the most susceptible to stall, responding

with the highest stress levels. The moderately swept SR-3 and the highly

swept SR-5 models remained stable at increasingly higher blade angles and

RPM's than the SR-2, and also responded with lower stresses. As expected,

all three models e.ncountered high stressing at the highest blade angles and

rotatlonal speeds. It is believed that these were forced excitation responses due to vortex shedding, or buffeting.

The test data show that the strain gages were properly located to allow the

various blade vibratory modes to be distinguished. Data analysis indicates

that stall flutter responses occur in the third mode (torsion) for the SR-2

model and in the second mode for the SR-3 and SR-5.

Vibratory blade stresses measured during a similar independent test conducted

in the NASA/Lewls lOxIO wind tunnel show very good agreement with the UTRC

test data.

Stall flutter calculations were made using a recently developed flutter analysls method that can determine the stability of thin, highly swept blades,

such as those used on Prop-Fans. The onset of stall flutter is analytically

determined to be at operating conditions for whlch blade damping goes to

zero. Negative damplng indicates an unstable condition.

Flutter predictions for the three Prop-Fan models were made and compared to

test data. Flutter boundaries were determlned from the test data, based on

the occurrence of steeply rlslng stresses with increasing blade angle or rotor RPM, since damping was not measured. The calculations show negative

damplng occurring at generally the same operating conditions for which high

stresses were encountered during test. Very good agreement was seen for the

SR-2 and SR-3 models, with less agreement for the SR-5 model which did not

give strong flutter indications dur!ng test. However, the tested trend showing stability to increase with blade sweep was well predicted. The theory

predicts that stall flutter will occur in the third mode for all three models. This agrees wlth the SR-2 test data, but not with SR-3 and SR-5 measurements.

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SYMBOLS

Blade Actlvlty Factor

I00,000

Blade Sectlon Chord N1dth, m

Blade Sectlon Deslgn Lift Coefflclent

Power Coefficient = SHP/p n2D 5

Rotor Diameter, m

Rotor Speed, RPM

Rotor Speed, revolutlons/sec

Rotor Torque, N-m

Shaft Horsepower

Non-Dimensional Blade Radius

Reference Blade Angle, deg

Blade Angle at 3/4 Radlus, deg

Air Denslty, kg/m 3

1.0

bx3dx

O

0.2

Sl units of measurement used throughout unless speclfied otherwise.

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!
1.0 INTRODUCTION

The occurrence of fuel shortages, increased fuel cost and the threat of future worsening conditions for alr transportation has caused NASA to sponsor

studies of new, more efficlent, aircraft and propulslon systems. One of the

promising concepts estabTished by these studies is the advanced high speed

turboprop, or Prop-Fan. This propulsion system differs from existing turboprops. The Prop-Fan has greater solidlty than a turboprop, achleved by more

blades of lar_er chord. The turboprop has straight blades with relatively

thick alrfoll sections; the Prop-Fan has swept back blades with thin airfoil

sections to enhance performance and reduce noise. The turboprop cruises at

no more than 0.65 Math number; the Prop-Fan is designed to cruise at 0.7 to

0.8 Mach number. The diameter of the Prop-Fan Is about 40 to 50% smaller

than that of the turboprop. For maximum performance the Prop-Fan makes use

of advanced core engines of the kind being used in modern turbofan engines.

Performance Is also enhanced by use of a spinner and nacelle aerodynamlcally

contoured to reduce compressibility losses by retarding the high velocity

flow through the root sections of the Prop-Fan blades.

Utilizing predicted and measured aerodynamic performance data, weight estlmates, and noise projections, several Government sponsored studies by both

engine and airframe manufacturers have concluded that a fuel savings of approximately 20 to 40% dependlng on operating Mach number should be achleved

by a Prop-Fan aircraft, as compared with a high bypass ratlo turbofan alrcraft. With these encouraging results, a research technology effort has been

instituted to establlsh the deslgn criteria for this new propulsion system.

A major objective In the development of Prop-Fan configurations Is to insure

the structural integrity of the rotor. Slnce the Prop-Fan is such a significant departure from conventional propellers, with its highly swept, thln

blades, the structural demands are substantial. The high speed operatlon of

highly swept blades imparts large forces to the llmlted materlal Inherent to

the thin airfoil sections needed for efficlent performance. It is imperative

that the rotor be able to absorb the aerodynamic loads at all operatlng condltlons, as well as the centrifugal loads associated wlth its unique shape

and construction. The steady-state dynamic response of the blades must be

low and flutter instabilities must be avoided, for safe operation.

As part of the continuing studies of Prop-Fan structural stability being conducted by Hamilton Standard, under contract to NASA-Lewls Research Center,

static stall flutter tests were conducted on the SR-2 8-bladed, SR-3

8-bladed, and SR-5 lO-bladed model Prop-Fan conflguratlons. These tests were

conducted during September and October, 1981 at the United Technologies

Research Center.

Thls report summarizes the results of this static stabllity investigation.

Included are trends of the measured blade stress test data with operating

conditions for the three models. Blade vibratory stress data were analyzed

for the peak stress amplltudes of the total signal as well as for the

frequencies and amplitudes of the spectra] components. In addition, stall

flutter stability boundaries were predicted using a theoretically based calculation procedure for comparison to test results. The comparisons were used

to evaluate the accuracy of the prediction methods and to recommend improvements to increase their effectiveness as Prop-Fan design tools.

1/2

2.0 DESCRIPTION OF EXPERIMENTAL PROGRAM

2.1 Model Description

Three Prop-Fan models were Installed on an isolated nacelle In the United

Technologies Research Center wlnd tunnel and were tested to determine the

dynamic stabil|ty in stall (see Ref. l). The models were designated SR-2,

SR-3 and SR-5 and are shown installed In Figures l through 3, respectlvely.

The blades are made wlth a solld metal construction, and the planforms are

characterized by increasing sweep, from the straight bladed SR-2 model

Prop-Fan to the highly swept SR-5 model Prop-Fan. Figure 4 is a schematic

showing these planforms along wlth strain gage locations which wlll be

discussed later.

The SR-2 is an elght-bladed model constructed of steel. The SR-3 Is an

elght-bladed model and the SR-5 is a ten-bladed model, both of which were

constructed of tltanlum. Table I shows some of the deslgn parameters for

these conflguratlons. All of these configurations are derived from full

scale designs that are intended to operate at a rotational tlp speed of 800

ft./sec, and at 0.8 Mach number f11ght speed. Figure 5 shows the variations

of many of the geometric parameters of each deslgn.

2.2 Test Models

Each of the three test models comprised an approximate I/8 scale, variable

pitch (ground adjustable), 62 cm (24.5 in.) diameter Prop-Fan configuration.

Each model consisted of a unique hub, blades, and splnner as well as a common

nacelle afterbody. The blades, hub, and spinner were designed and fabricated

by Hamilton Standard. The nacelle afterbody was fabricated by UTRC per

Hamllton Standard design. Each model was deslgned for counterclockwise rotatlon (vlewlng upstream).

The blade roots were equipped with a gear-sector that engaged a common ring

gear in the hub, which assured blade pitch angle synchronlzatlon and slmpIified blade angle changes. The gear-sectlon mechanlsm permits an infinite adjustment In blade angle over approximately a 90 degree range. However, a

locking pln, which is inserted in the ring gear and Indexing plate holes,

results in incremental settings of l degree. The maximum blade pitch angle

settings for all three models was limited to 80 degrees. The minimum setting

varied for each of the three models and was limited by mechanical

interference at the blade roots to -14.3 degrees for the SR-2, -8 degrees for

the SR-3, and +ll degrees for the SR-5 model. Blade pitch angle was measured

by placing the partlcular blade in a horizontal posltlon and employing an

inclinometer fixture on the face side of the blade at 0.78 radius, known as

the reference station. Blade pitch angle Is defined as the acute angle

between the blade chord and the plane of rotation. Prior to installation,

each model rotor was statically balanced on knife edges, and materlal was

removed from the heavy side of the hub by drilllng holes to provide a static

balance.
2.3 Wind Tunnel Faciltty

The United Technologles Research Center (UTRC) Large Subsonic Wind Tunnel

(LWST) shown in Figure 6 is a single-return, closed-throat facility with Interchangeable 5.5 and 2.4 meter (18 and 8-ft.) octagonal test sections. Maximum tunnel velocity is approximately 90 m/s (200 MPH) in the 5.5 m (18-ft.)

test section and near sonic Mach numbers can be obtalned in the 2.4 m (8-ft.)

test section.

For the subject statlc test program, the tunnel circuit was arranged (Figure

7) to reduce tunnel wall effects and to mlnlmlze reclrculatlng flow through

the plane of'the propeller. This was accomplished by locatlng the 5.5 m

(I8-ft.) diffuser In its normally stowed position, thus permitting unobstructed airflow to enter the downstream end of the test section. Flow recirculatlon through the tunnel circuit was mlnlmlzed by blocking the open circuit

which normally mates to the diffuser and by exhaustlng the propeller airflow

through the air exchanger valves which were set In the l m (3-ft.) open pos-

|tion.

The LSWT has available both static and dynamic data acquisition and recording

systems. This test program used the static system called Online Computer

Controlled Acquisition Recording (ONCOAR). Its mlnlcomputer inltialized and

controlled that data acquisition equipment, acquired data, displayed and recorded the acquired data, and transmitted the data via a Multi-Serlal Transmission (MST) line to another high speed digital minicomputer system for onllne processing. The reduced data were then displayed in tabular form on a

computer terminal or In graphical form on a cathode ray tube. ONCOAR is capable of acquiring analog data on up to 25 dffferent channels, using up to

eight scanlvalve or temperature scanner solenoids. In addition, the system

was set up to accept input from up to 14 digital channels. Thls list

Includes six channels for the main balance, one each for model pitch and yaw

attitudes, barometric pressure, test section pressure dlfferential, tunnel

stagnation temperature, two channels for Events Per Unit Time (EPUT) signals

and one for a precision pressure transducer/regulator. ONCOAR is capable of

recording and storing up to 1200 pieces of analog or digital data in any

combination within the above llmits.

In this test, ONCOAR recorded a total of 219 pieces of data per polnt on nine

analog channels and four digital channels. Approximately 30 seconds were

required to acquire the data, and an additlonal I0-15 seconds were needed

(depending on computer workload) to transmit, reduce, and display the on-line

data for a total of approximately 40-45 seconds per data point. The raw

data, which had been recorded on floppy disc by ONCOAR, were transferred to a

nlne-track magnetic tape in large computer compatible format for further

processlng off-line. This off-11ne processing can be used for correcting

data as well as for refining processing procedures.

A dynamic data recording system supplled by Hamilton Standard was used to

monitor and acquire time variant blade stress data. This system provides

eight channels of signal condltionlng and amplification, FM recording and

playback capability, oscilloscope monitor, and switching gear to acquire up

to 16 channels of strain gage type data.
2.4 Propeller Dynamometer

The Prop-Fan model was driven by the UTRC Prop-Fan test rig dynamometer

(PTR). It uses two varlable-speed motors housed withln a streamline caststeel pod with an integral support strut (Figure B). The motors are mounted

In hydrostatlc bearlngs to restrain all motion except axial and rotational

motion about the longitudinal axis of the dynamometer. These motions are restrained by load cells which measure thrust and torque of the mode! PropFan. Each motor has a nominal rating of 280 kW (375 hp) at 12,000 RPM; together they provide a maximum torque of up to 450 N-m (330 Ib-ft.) over the

entire speed range. Model speed is controlled by variable frequency power

supplied by two motor generator sets and measured with an events per unit

tlme meter and a 60-tooth gear signal generator. Prop-Fan rotational dlrection for this test was counterclockwise looking Upstream. The dynamometer is

falred such that there is a minimal axial static pressure gradient through

the plane of the Prop-Fan and so that the Prop-Fan rotor and splnner surfaces

are the only portions of the metric system exposed to the alrstream. Pressure Instrumentation is provided withln the dynamometer to correct measured

thrust for any differential pressure between the front face of the hub and an

equal area in the rear fairing.

The Prop-Fan dynamometer was mounted on the floor at the downstream end of

the 5.5 m (I8-ft.) test section (facing south). This posltloned the models

within 25 cm (lO-in.) of the open tunnel circuit (Figure 9). The Prop-Fan

drew air from the courtyard in an area unconfined by tunnel walls and discharged it Into the tunnel circuit. With the tunnel circuit blocked at the

extreme south end of the courtyard and the alr exchanger vaTves open approxlmately l m (3-ft.), the flow created by the Prop-Fan passed out the a|r exchanger valves and could not reclrculate through the plane of the propeller.

The relationship of the dynamometer, test section, courtyard, and blocked off

tunnel circuit is shown in Figures lO and If.

Dynamometer Instrumentation conslsted of: thrust and torque load cells, a

I/rev reference slgnal, a 60/rev slgnal for RPM, vertical and lateral plane

vibration transducers, bearing and motor thermocouples, and internal cavity

pressure taps. The instrumentatlon electrlcal and pneumatic lines were routed down through the hollow PTR pylon to the tunnel floor and from there to

appropriate monitoring and recording devices in the control room.

2.5 Model Instrumentatlon

Each of the three test model Prop-Fans was instrumented with straln gages on

the camber surface to measure bending and torslonal stresses on four blades.

The strain gages were located at the maximum principle stress locations of

the natural modes, as determlned by analysls. The locations of these straln

gages for each blade model are documented in Figure 4. The blade strain gage

configuration for each of the three rotors Is described in Table II. Blades

were numbered sequentlally around the rotor in a clockwise directlon when

viewed from the rear. Blade strain gages are Identified by BGx-y, where BG

designates blade gage, x Is the blade number and y is the gage number. The
electrical lead wires were routed from the strain gages along the trailing

edgesof the blades and through the hub to a slip ring assemblymountedon

the upstream surface of the hub. An electronic, two-positlon switch on the

rotating portion of the slip ring assembly permitted the selection of either

of two groups of five strain gages to be monitored. The electrical leads

from the stationary portion of the slip ring assembly were routed out the

front end of the spinner (Figures 2 and 3) through a pneumatic air cooling

llne and from there to the approprlate HSDmonitoring equipment in the tunnel

control room. Air cooling was provided to each of the eight rotating

elements of the sllp ring through a 1.3 cm (0.5-1n.) diameter tube connected

to a 138 kPa (20 pslg) filtered alr supply.

A static pressure probe was mounted In the plane of the airflow entrance to

the 5.5 m (18-ft.) test section approximately 218 cm (86-In.) radially from

the prop centerllne (Figures 9, lO and ll) to pcovlde an indication of tunnel

through flow as a result of propeller thrust. This probe was connected to a

high accuracy, low pressure transducer, SETRA 140 Pa (0.02 psig) capability,

which provided wind gust data to the dynamic data system, as well as steadsstate data to the static ONCOAR data system. In addition, a tunnel spannlng

pressure rake was mounted 109 cm (43-In.) starboard of the prop centerline

(Figures 9, lO and ll) to provide steady-state wind speed data. The 13 elements on thls rake were routed to a water manometer board In the tunnel control room. However, due to the low velocities, and hence low pressures, thls

system could not provide the desired resolution. For most of the tests, local tunnel velocity was measured solely by the static probe/SETRA system.

Also, a conventional, vertical axis, cup anemometer was used for visual reference of the ambient wind condition <Figure IO).

2.6 Test Procedures

The primary objective of the test program was to define the stall-flutter

boundaries, if any, of the SR-2, SR-3, and SR-5 Prop-Fan models under static

Flow conditions. This was accomplished by conducting rotational speed sweeps

from 2000 RPM to maximum and back to 2000 RPM, at Fixed blade pitch angles,

while continuously monitoring blade stresses and recording these stresses on

FM tape. Performance data, including rotor thrust, torque, total pressure

rise, and nacelle surface pressure distribution, were acquired at regular,

discrete rotor speed intervals.

Typically, a test run was conducted as follows: the blade pitch angle, at

0.78 radius, was set using the appropriate fixture and inclinometer; water

cooling, oll lubrication, and hydrostatic pressure and scavenge systems were

activated; a start zero was acquired on both ONCOAR and on the FM system; the

rotor was brought on-llne at a rotor speed of approximately 2000 RPM; all ten

strain gages were monitored prior to rotor acceleration; rotor speed was

increased from 2000 RPM to maximum in a slow, continuous sweep while blade

stresses were monitored.
The rotor speed sweep was restricted by blade stress limits which differed

for each of the three models. In addition, a speed sweep could be limited by

the maximum available electric rotor torque. For thls program, thls appeared

to be approximately 410 N-m (300 ft-lb.). The ultlmate llmlt In rotor speed

if stresses and power permitted was 9000 rpm, whlch corresponds to a rotor

tip speed of approximately 293 m/s (960 Fps). Steady-state and dynamic data

were recorded at the maximum rotational speed and then in increments of 500

RPM between the maximum speed and 2000 RPM. Thls procedure was repeated at

different blade pitch angles and model configurations For approximately 55

data runs. The conditions at which the three models were tested are summarized in Table III. Also summarized in Table III are the conditions for

which calculations were performed and will be discussed later.

Since the test program was conducted under static (no flow) conditions,

thrust and torque tare data were not acquired nor applied to the actual performance data.

2.7 Operating Conditions

The operatlng conditions used ?or the calculations cover a large range of

blade angle settings and rotational speed settings. These conditlons are

presented in Table III for the test runs as well as those for the computations. Since the calculations involve the use of the lengthy MSC NASTRAN

program for the mode shapes and Frequencles, the number of runs was mlnimlzed

in order to reduce the computer usage. The MSC NASTRAN program was therefore

run at blade angles of -lO ° and 55°. The Frequencies used in the stability

analysis were Interpolated For conditions wlth blade angles other than those

calculated using MSC NASTRAN. The mode shapes of the nearest MSC NASTRAN

case were used For the stability analysls.

The blade angle schedules in Table III For the static test conditions are

different For each of the models. The RPM schedule is the same For each model except that the upper llmit Is restricted by either a power limit or a

stress limit. All test polnts and calculations were at sea level conditions.

2.8 Data Reduction

Blade vibratory stress data were displayed and monitored, on-line, on a

multichannel oscilloscope. Hamilton Standard personnel interpreted these

time-varlant data In a continuous, on-line manner throughout the test program. Test condltlons were selected and operational limits were observed as

a result of this (on-line) monitoring. In addltlon, stress data, For each

steady state data point and a11 rotor acceleratlons, were recorded on FM

tapes which were retained For comprehensive, detailed analysls.

The analog tapes were analyzed by obtaining total vibratory stress amplitudes

using electronic peak stress converters and recordlng the resulting slgnals on

strip charts. As a second step, samples 30 seconds In length from the magnetic tape were processed using a real tlme analyzer. These samples were time

averaged to produce spectral analyses of the data. This information, in turn,

was then stored on tape for a permanent record of each case. The data were

then transmitted to a high speed digital mlnl-computer for processing. At

this point, a computer program was used to pick out the peak amplitudes and

the associated frequencles. These were then tabulated and printed according

to case number and condition. Automatic routlnes were developed that produce

Campbell diagrams and vibratory stress vs. RPM For each blade angle. These

items are discussed Further In the spectral analysls section (see 4.3) of

this report.

7/_

3.0 ANALYTICAL TECHNIQUES

3.1 Method Descrlption

The method used to estimate stall flutter boundaries involves several parts.

The various computer programs used are listed by designation and purpose in

Table IV. The primary analysis used to calculate these boundarles is the

F203 analysls, and the other programs are used to generate data for or from

this analysis.

The F203 stabillty analysls was developed by J. Turnberg (Reference 2) prlmarlly for classical flutter. It is a linear elgen-_alue solutlon that uses

unsteady aerodynamics accounting for compresslbillty effects and blade

sweep. For classical flutter, the quasl-steady llft analysls uses the value

of 2_ for the lift curve slope. The computer program has a separate portlon for stalled conditions that is used to calculate stall flutter. Here

the unsteady aerodynamic analysis uses a parabolic pressure dlstrlbutlon for

determining the unsteady forces. For the quasl-steady terms In stall, the

program uses the lift curve slope at the local angle of attack for a particular operating condition. In addition, the analysis uses a method developed

In Reference 3 for stalled flow. This method complements the eigen-solutlon

and gives results that are very slmllar. It uses an energy balance that relates the energy developed by the aerodynamic forces to the straln energy in

order to determine the damplng of the system. It employs the same unsteady

aerodynamic terms as are used in the stall flutter elgen-solution.

The stability analysis F203 is also a modal analysis that requires threedlmensional modes, developed in the blade chord coordinate systems, at each

blade spanwlse station. Generally, other linear aeroelastic analyses de'

scribing rotating aeroelastic surfaces will approximate the geometric blade

angle relative to the plane of rotatlon using small angle assumptlons. In

static operation, this angle is very large, up to 70 degrees for a Prop-Fan.

The F203 analysis uses the blade chord as a coordinate system such that the

small angle is made on the section angle-of-attack, which is small for most

applications. The input requlres that the mode-shapes and modal masses be

transformed to the above mentioned coordinate system. Generally, the mode

shapes and frequencles are developed by the beam analyses, H025 and H027, or

the finite element methods, NASTRAN or BESTRAN. A program called F214 makes

the necessary transformatlons from finlte element methods while approximating

the blade motions by three-dlmenslonal beam type displacements. Chordwlse

deformations are approximated by a rigid section. The methods used In the

present analysis are discussed In more detail in Reference 2.

Figure 12 shows a block diagram of the procedure used In the stall flutter

analysis. It can be seen in thls dlagram that the output from the finite element methods are Input for the F214 coordinate transformation program. (It

should be mentioned that there Is an earller modification to the F.E. data by

a program called "MODES". This rotates the data for each element Into the

shaft plane and modifies the format. It is not shown on the block dlagram.)

The operation of the F214 program can be implemented by the CLIST Control

Program as shown by the block diagram in Figure 12. An output file from F214

is created containing the transformed mode shapes, modal masses and frequencies.
The aerodynamicproperties used for the F203stabillty analysis are initiated

in a data bank accessed by the H444performance analysis, where the data for

several airfoil shapesare stored. Oncethe performance has been determined

at the operating condition of interest, the lift and momentslopes are then

determined as a function of angle of attack at each radial station for this

operating condition. These slopes determine the unsteady and quasi-steady

loads In the stablllty analysis.

As shownIn Figure 12, the runnlng of the F203stability module is controlled

by the F203CLCLIST. Here the transmission of the input and output files is

managed,and the plot program is executed. The plot program PLT203was

created to run from a file that consists of data for manyF203 runs. The results of thls programare plots of the printed output, where dampingand frequency are plotted as functions of blade angle.

It is suggested here that the stall flutter boundaries predicted by this

analysis maybe conservative. Th|s is partially due to the fact that stall

flutter is a limit amplitude phenomenon,and can exist at small amplitudes.

If the limit amplitude is small enough, then it is posslble that flutter will

not be noticed experimentally, because It will be lost in stresses due to

turbulence or other causes. The present analysis Is a linear analysis and

can, therefore, predict only the onset of flutter, which could be at low

stress levels. Thus, the predicted boundary would appear conservative, in

relation to the point of measuredhigh stresses.

3.2 Calculated Instabilities

Calculations to estimate stability boundaries were made for the SR-2, SR-3

and SR-5 model blades using the F203 stability analysis. Values of total

damping were calculated for 5000, ?000 and 9000 RPM, at many blade angles, as

shown in Table Ill. Flgures 13 and 14 show the damping to critical damplng

ratlo for all three models as a function of blade angle at 7000 and 9000 RPM,

respectively. The onset of stall flutter is assumed to be at the point where

the damping goes through zero. At 7000 RPM, it is seen that increasing the

sweep is beneficial in delaying the stall flutter to a higher blade angle.

Note that the SR-5 does not flutter at 7000 RPM but is delayed until 9000

RPM. All of the stall flutter predictions are third mode instabilities. No

instabilities were calculated for the first or second modes. Figure 15 shows

the typical damplng ratio relationships between the modes for the SR-3 model

Prop-Fan blade. The flutter boundaries, as functions of blade angle and RPM,

will be shown later in discusslons of the test results.

T

10
4.0 TEST DATA EVALUATION AND COMPARISON WITH CALCULATIONS

4.l Response Frequencies

The calculated blade response Frequencies for the SR-2, SR-3 and SR-5 model

blades are shown In Figure 16, where blade frequency Is plotted as a function

of rotational speed. Also shown in this Figure are data points taken from

spectral analyses of the analog blade stress data, some of which will be discussed later. As previously mentioned, the SR-2 blade frequencies were calculated using the beam methods H025 and H027, while the SR-3 and SR-5 blade

Frequencies were calculated using the MSC NASTRAN analysis.

It is seen that good agreement exists In all modes between the test results

and the computations for the SR-2. Note that the slopes of the second and

third mode show good agreement. For the swept models, the SR-3 and SR-5,

good correlation Is made for the first two modes wlth poorer correlation occurring for the third, fourth and Fifth modes. However, good agreement is

seen for the slopes of the higher modes for these two models. Both swept

blade models show a measured response between the second and thlrd calculated

mode. The nature of thls response Is not underst_d at thls tlme.

4.2 Total Stress Results

As prevlously indicated, total peak vibratory stress was recorded on strip

charts for the SR-2, SR-3 and SR-5 model Prop-Fan blades. The stress data

from those charts were tabulated and selected data were plotted on curves of

total vibratory stress (infrequently repeating peak stress*) as a function of

RPM for various blade angles. These p}ots are shown In Appendix A For the

three Prop-Fan models. Additionally, cross plots were made to produce stress

contour plots for the n_de] Prop-Fan blades. These are contours of constant

total vibratory stress, plotted on curves of reference blade angle vs. RPM,

and are shown In Figure ]7 through 19. Note that the takeoff design operatlng point for each blade is shown for reference.

Isostress Contour Plots - Figure 17 shows the total stress contours for the

tip bending gage and the shear gage outputs of the SR-2 model <blade number

5). Both gages show the highest stre_s at a reference blade angle of 40 degrees and 7000 RPM. From Figure 16, It is seen that thls Is very close to

the third mode 5P critical speed. The buildup seems gradual with Increaslng

RPM and less gradual with Increasing blade angle. This effect Is probably

due to the fact that a change in reference blade angle has a greater effect

on the blade angle of attack than a change in RPM. These results are typical

for conventional propellers that encounter high stresses in the statlc conditlon. Since the third mode is the torsion mode, it Is not surprising that

stall conditions combined with critical speed effects would cause a stress

buildup. It Is also noted that the gradual buildup makes It difficult to

find a precise definition of a stall flutter boundary, especially one where

the damping might be considered as having a value of zero.

*The Infrequently repeating peak Is defined as the maximum stress peak that

repeats itself two or three times during the stress data sample period.

11
The calculated flutter boundaries for the SR-2are also shownIn Figure 17.

The calculated boundaries represent the torslon (third) modewhile the measured total stress represents all the modes. A spectral methodby which the

modal stresses can be separated will be discussed in the next section.

Similar Isostress contour plots are presented In Figure 18 for the SR-3 model

Prop-Fan. This figure represents the output from the Inboard bending, the

shear and the tip bending gages, respectlvely. In order to smooth out some

of the Irregularities In the data, the values of stress were averaged between

blades I, 2, 5 and 6 for the shear and tlp bending gages, and between blades

I and 5 for the inboard bending gage. These curves show three entirely different patterns. For example, the shear gage shows a very gradual Increase

In stress with Varying RPM and blade angle. However, the Inboard bending

gage shows a sharp Increase in stress near 40 degrees blade angle and 6000

RPM. The tip bending gage indicates a sharp rise in stress near 30 degrees

blade angle, but shows a gradual increase with RPM.

Subsequent viewing of osci11ograph records clearly shows different predominant frequencies of similar amplltude occurring on different gages of the

same blade for some records. This indicates that stalled flow can excite

several different modes slmultaneously. It can also be concluded that the

strain gages were effectlvely placed to measure the response of each mode.

Interestingly, the flutter indications predicted for the shear gage seem to

occur experlmentally for the tip bending gage. Spectral studies made for the

SR-3, and discussed later In this report, shed more light on this apparent

discrepancy.

Figure 19 shows stress contour plots for the SR-5 model blade. They represent the output from the inboard bending gage and the shear gage, respectively. The shear gage shows a very high stress peak at 6500 RPM. This can be

attributed to the fact that it Is very close to a 6 per revolution critical

speed for the 4th experlmental mode, as shown In the Campbell diagram in

Figure 16. This mode colncldes with the 3rd predicted mode. High 4th mode

response Is also indicated on spectral plots, to be shown later in the report.

The calculated stall flutter boundary predictions are also shown in

Figure 19. These were developed for the 3rd mode and represent the boundary

of zero damping. It Is seen that this predicted boundary occurs at very high

RPM and does not coincide with any sudden stress rlse. Some of the lack of

correlation between test and prediction might be due to the fact that the

test results Include aerodynamic excitation other than stall flutter, such as

buffeting. Also, it may be difficult, In some cases, to distinguish between

stall flutter response and a critlcal speed crossover.

12
Stall Flutter vs. BufFet - It may be useful to discuss the differences between stall flutter and buffeting. Buffeting is defined as a for_ed excitation due to an instability of the air, such as vortex sheddlng, shock oscillation, or turbulence. Stall flutter Is an instability due to the Interaction between the alr and the blade. In stall Flutter conditions, the motion

of the blade and the aerodynamic loading on the blade are strongly interdependent. In buffet conditions, the motion of the blade has little effect on

the loading.

Generally, as blade angle is increased the Prop-Fan progresses From normal

load conditions to stall and then to deep stall. Stall flutter can occur as

the Prop_Fan becomes stalled and buffet occurs in deep stall. At a specific

operating condition, the local angles Of attack along to blade span increase

as the blade angle is increased, with stall first occurring inboard and then

progressing outboard.

In order to define when the Prop-Fan Is stalled, the blade reference station

(0.78 radius) Is generally a good point to consider as being a stall control

station. The conditions at which the current Prop-Fan blades stall was not

investigated for this analysls. However, from preliminary estimates it Is

thought that stall occurs at a reference blade angle between 30 and 35

degrees, For Prop-Fans at statlc (zero Forward speed) conditions. Although

the boundary between stall flutter and buffet regions is not clear, it Is

thought that buffet occurs at blades angles which are substantially higher

than blade angles for which stall flutter occurs. For thls discussion, the

buffet region Is defined to be at blade angles of approximately 45 degrees

and larger.

Blade Stall vs. Rotor Torque - Prop-Fan rotor torque can be an Indicatlon of

the loading condition on the blades. F_gure 20 shows the measured shaft

torque, as a function of reference blade angle for various RPM, for the SR-2,

SR-3, and SR-5 model Prop-Fan configurations. Each plot shows a variation In

RPM From 5500 to 8500 RPM.

Generally the torque increases with blade angle and RPM for a11 configurations. It is seen from these curves that there is a change In the torque at

or near the blade angles where stall might be expected. The SR-2 shows the

greatest effect, where the torque increases rapidly near a blade angle of 28

degrees, peaking at 30 degrees and returning to the torque curve at 33

degrees.

The SR-3 data show a decrease in torque near a blade angle of 31.5 degrees.

It Is not known if there Is a torque rlse just before this polnt because of

Insufficient data. The change in torque seems less severe than that observed

for the SR-2.

The SR-5 data show a small depression at a blade angle of 34 degrees for the

higher RPMs. This Is a lesser effect than that seen for the SR-3. The low

RPM SR-5 data show llttle of thls effect.

13
The effect of stall on the torque curves is most severe for the SR-2 and

least severe on the SR-5 with the SR-3 falllng in between.. This indicates

that the Influence of stall on the torque is affected by blade sweep, since

the major difference between the conflguratlons is sweep, the SR-2 being

non-swept and the SR-3 and SR-5 having increasing sweep, respectively.

It is also noted that the torque change occurs at an earlier blade angle on

the SR-2 and progressively later on the SR-3 and SR-5, respectively. The

test data discussed above (Figures 17 to 19) show that a high stress rise

occurs at blade angles near where the torque inflections occurred. Also, the

highest stresses occurred on the SR-2, with progresslvely lower stresses on

.the SR-3 and SR-5. Thls indicates that stall and/or stall flutter occurs at

similar conditions as the inflections on the torque curves. It is therefore

concluded that the torque curves can indicate the presence of stall or stall

flutter condltlons.

It is recommended that in future static tests on Prop-Fan models, fine varlations be made in RPM and blade angle in the area Just below, in and above the

stall condition, and that torque measurements be made at each steady state

condition. This would be helpful in defining the condition of blade stall

onset and its relatlon to blade stress.

Blade Stress vs. Damping - It should be noted that some of the difficulty, in

comparing calculated stall flutter boundaries to the experlmental results, is

due to the nature of the parameters which are used to define the boundary for

each. The calculated stall flutter boundarles are linearly determined to be

at the point where the critlcal damping ratio goes zero. The experimental

flutter point is determined to be where there is a sudden rlse in vibratory

stress with increaslng RPM, usually to a high stress value. This Ignores the

fact that in a non-linear system, the damping can go to zero at flutter onset

but can also be zero at some llmit amplitude. It Is conceivable that the

limit amplitude could be small, while the damping Is zero. It may be misleading to investigate stall flutter condltlons by comparing the two different parameters of damping and stress, as was done here. A better result may

be expected if a non-linear aeroelastlc analysis is used to produce stress

predictions that could be compared to the experimental stresses. At the time

of this work, however, a reliable analysis of this nature was not available.

4.3 Spectral Analyses

Measurements of total stress cannot be used to fully characterize blade

dynamic behavior. For example, total stress values do not allow the stress

contributions of each mode or P-order response to be distinguished. Spectral informatlon is helpful in evaluating modal stresses. This is examined

in the form of spectral plots of vibratory stress as a function of frequency.
SR-2 Results - Figure 21 Is a spectra] plot of measured stress For the blade

tip bending gage output on the SR-2 model operating at 7000 RPM and a reference blade angle of 36.2 degrees. Figure 22 Is a spectral stress plot of the

shear gage at 8500 RPM and a reference blade angle of 31.5 degrees. These

two Figures represent condltlons In the high stress areas For each gage, as

seen in Figure 17. They are not the conditions of highest stress, but are

located In the area of steep stress rlse.

The Indications from Figures 21 and 22 are that, for the SR-2, the flutter

occurs In the third mode at or near 600 Hz. Thls mode Is considered the primary torslonal frequency (See Figure 16). The th|rd mode response level seen

In Figure 21 Is large due to its proximity to the 5P critical speed. Figure

22 shows substantial twlce per revolution response. This is unexplained,

except that it Is a relatively low stress, and this condition may be close

enough to the 2P critical speed to give some magnification to the 2P stress.

Figure 23 shows a spectral plot for the mld-blade bending stress at 5000 RPM

and reference blade angle of 50.3 degrees. Here the response Is substantially In the first mode. This may be a buffet condition excitlng the first

mode wlth some 2P magnlfication due to the nearness of the 2P crltica] speed

(See Figure 16).

SR-3 Results - Spectral plots from SR-3 testing are shown In Figures 24, 25

and 26. Tlp bending stress Is shown In Figure 24 for a condition of 9050 RPM

and 31.7 degrees reference blade angle. This conditlon Is In a steep stress

rlse area (See Figure 18) that Is indicative of stall flutter. Figure 24

shows the tip bending to have a high 3P response accompanied by a moderate

second mode contribution. The 3P response seems exaggerated by low damping

assoclated with the 2ridmode response. The 4P, 5P and 6P responses cou]d

also be critical speed related (See Figure 16).

A more clear example of sta]1 Flutter response Is shown In Figure 25. The

tip bending gage spectrum In thls figure Is for a condition of 32.7 degrees

blade angle and 7020 RPM. This Is not near any crltlcal speed and is also in

the steep stress rise area. Figure 25 shows substantial second mode response

with no apparent excitation. Thls isa strong indication of stall Flutter

response. There Is also response present in the third, Fourth and Fifth experlmental modes. Recall From F|gure 16 that what is termed the Fourth experlmental mode is shown near the third predicted mode. This mode shows the

least response In Figure 25, which is contrary to the stall flutter predictions discussed earller.

Figure 26 represents a 5010 RPM and 50.3 degree blade angle condit_on. This

is considered to be In a hlgh stress buffet reglon due to the large blade

angle, as d_scussed earlier. Th_s Is probably not a stall Flutter condltion. Flgure 26 shows primarily Ist mode response. The contour plots In

Flgure 18 also show mostly Inboard and tip bending at this conditlon. Spectral plots of the shear gage signals (not shown) Indicate comparatively

little stress.

15
SR-5 Results - Two SR-5 spectral stress samples are shown in Figures 27 and

28. The first represents the output of a shear gage at 8500 RPM and a reference blade angle of 35.7 degrees, and the second is the output Of an inboard

bending gage at 6500 RPM and a reference blade angle of 49.8 degrees. All

the stress peaks shown for these two curves indicate relatively low stresses,

but the shear gage seems to be responding to whlte noise type excitations,

This indicates the possibility of buffetlng, and there seems to be no evidence of a self excited response. This is also seen in Figure 19, In that

there is no sudden stress rise in either the shear or bending gage. The inboard bending spectral curve (Figure 28) shows a low level second mode response, and little of anything else.

The indications from these data are that the highly swept SR-5 Prop-Fan model

has little or no stall flutter problem, the SR-3 has a moderate stall flutter

response, while the SR-2 has a strong stall flutter response. Thus, sweep

seems to have a suppressing effect on stall flutter.

4.4 Modal Response for the SR-3

The stress peaks that were obtained from the spectral analysis and used in

the Campbell diagram of Figure 16 can be categorized as to frequency and

mode. Table V indicates the frequency range assumed for each mode, based on

the experimental responses. Plots of stress vs. RPM for various blade angles

can be made for each mode and each gage. Diagrams of constant vibratory

stress contours can be plotted from crossplots of these curves. For this

report, only the Isostress contour plots for the SR-3 model will be shown.

Figure 29 shows the modal isostress contour plots for the SR-3 model Prop-Fan

stall flutter tests at the UTRC. Here, the measured modal stress is plotted

as a function of rotatlona] speed and reference blade angle. Each mode is

shown for the particular gage that generally has the highest response. Only

the first five modes are shown in Figure 29; one contour plot for each. For

the first plot (lst mode), it is seen that the high stresses occur at a reference blade angle of about 50 degrees and at RPM's greater than 6000. This

is well above what is considered stall, possibly indicating that these stresses are due to buffeting, which involves mostly the 1st bending mode. Note

that the identifying gage is the inboard bending gage, which is most responsive to the first mode. Also shown on this curve are the operational limits

of the test. These limits were established by drlve power limits, RPM limits

and blade allowable total stress limits.

The second mode is characterized by high stress, probably under conditions

for which the blade first encounters strong stall over most of its span.

This indicates stall flutter responding in the second bending mode. These

data corroborate the spectral results, discussed earlier. Note that the high

stresses are found prlmarily in the tip gage. Also, the stress does not seem

to be related to a critical speed, whereas the high stress observed in the

flrst mode could indicate the 2P crossover; see Figure 16. The higher modes

(3, 4 and 5) show little response. It should be noted that the calculated

results indicate that stall flutter should occur in the third mode. This is

inconsistent with the test results which show high stress occurring in the

second mode.

16
4.5 Comparison with NASA-Lewls Tests

Low speed stall flutte'r tests were conducted at NASA-LeRC in the lO x lO wind

tunnel during October 198l, and are reported In Reference 4. Some of the

tests were run at static conditions with a small component of velocity due to

Induction in the tunnel. Assuming this effect Is negligible, the total vibratory stress results observed at the UTRC were compared with those obtained

at NASA-LeRC. These comparisons are shown in Figures 30 through 32, where

total vibratory stress is plotted as a function of rotational speed for var1ous blade angles.

Figure 30 shows the total blade vibratory stresses for the SR-2 model PropFan. Shown are the outputs from the mld-blade bending, the shear and the tip

bendlng gages for reference blade angles of approximately 32 degrees and 40

degrees. Generally, the test results at the UTRC give stresses that are

similar to those obtained at NASA, except near or at crltlcal speeds, where

the UTRC results show higher stresses. This may be due to the fact that, at

the UTRC the rotor was subject to the effects of turbulence due to weather

conditions, slnce the test was open to the atmosphere.

Figure 3l shows the results from the Inboard bending, the shear and the tip

bending gages of the SR-3 model Prop-Fan blade. The comparisons are made for

reference blade angles of approximately 32 degrees and 60 degrees for shear

and tip-bending, and approxlmately 32 degrees and 50 degrees for the Inboard

bending. Note that the vibratory stresses are lower for the SR-3 model than

for the SR-2 (Figure 30) due to the benefits of sweep. The correlation between the results from the NASA-Lewis tests and the results from the UTRC

tests, for the SR-3, Is also very good. For the SR-3 bending gages, the vibratory stresses obtained from UTRC are somewhat higher than the NASA measurements, which Is the opposite from the SR-2 results. However, the UTRC

results for the SR-3 indicate higher response near the critical speeds than

the NASA results, as also occurred for the SR-2 model.

Figure 32 shows the results from the inboard bending, the shear, and the

chordwise bending gages on the SR-5 model Prop-Fan. Shown are the results

for the approximate reference blade angles of 32 degrees and 50 degrees.

Again, the UTRC vibratory stress results are somewhat higher than the NASA

data but the correlation Is still very good. As for the other blade models,

the UTRC SR-5 tests show higher response in the critlcal speed regions probably due to higher turbulence.

17/18

S.O CONCLUSIONS

As a result of the test and analysis program summarized in this report, the

following conclusions were reached regarding the static stabllity of the SR-2

straight blade, the SR-3 moderately swept blade and the SR-5 highly swept

blade Prop-Fan models:

I. Increased sweep tends to suppress the hlgh blade stresses caused by

stall flutter and buffet.

2. Correlatlon between tested and predicted Campbell dlagram modal frequencies was excellent for the first and second modes for all blade models.

, Correlatlon between tested modal frequencies and beam method caIculatlons

for the SR-2 model, at the higher frequencies, was good. Finite element

method frequency modal calculations for the SR-3 and SR-5 models showed

less agreement with test data at the higher frequencies.

, Comparisons were made between measured stall flutter boundaries, based on

steeply rising stresses wlth RPM and blade angle, and calculated boundaries based on zero blade damping. Good agreement between test and prediction was indicated for the SR-2 and SR-3 .models, while less agreement

was seen for the SR-5 model, which did not give strong flutter indications during test.

. Tested stall flutter response for the SR-2 straight blade occurred In the

torsional third mode, as was predicted. Test data for the SR-3 and SR-5

swept blades show stall flutter response primarily In the second bending

mode while the calculated results predict that stall flutter should occur

in the torsional third mode.

, Modal Isostress contour data Indicate that stall flutter and buffet occur

in different operating regions, with buffet occurring for very high

blades angles.

, Total vibratory stresses measured at static conditlons at the UTRC were

compared to those obtained in the lO x lO wind tunnel at NASA-LeRC, for

the SR-2, SR-3 and SR-5 models. Both the absolute stress amplitudes and

the trends with varying RPM agree very well for these two independent

tests.

19/2o

6.0 RECOMMENDATIONS

l . Since it was shown that there was l_ttle difference between testing In

the wlnd tunnel at NASA-Lewls or testing in the atmosphere at UTRC, it is

suggested that future static tests can be conducted in a wind tunnel.

This will eliminate duplicatlon of rig setup.

° It is recommended that in future static tests on Prop-Fan models, fine

variations be made In RPM and blade angle In the area Just below, In and

above the stall condition, and that torque measurements be made at each

steady state condition. This would be helpful in defining the condition

of blade stall onset and its relation to blade stress.

, The tests reported herein show varlations of measured blade stress wlth

rotatlonal speed (RPM) and with blade angle. It was observed that at or

near critical speeds, the testlng was limited to those RPM's for which

the stresses were below the limits. If the test condition envelope was

increased to include rotational speeds beyond these critical speed areas,

the scope of the data could be Increased. Additional understanding of

the phenomena of stall flutter and buffeting would develop if this could

be achieved.

, The correlation between the current stall flutter theoretical predictlons

and the experimental results can be improved. Deficlencles in the analysis may be due to its ]inearity. The analysis is linear in both the

aerodynamlcs and structural dynamics by assumlng small amplitude displacements. The actual blade response in stall flutter very often has

large amplitude displacements. This behavior requires that non-linear

aerodynamlcs as well as non-linear structural response be included in the

analysis for proper representation. Also, Corlolls forces due to rotatlon are non-llnear for large amplitude vlbratlons.

It is recommended that a non-linear analysis be developed that can model

the behavlor described above. It is suggested that this analysis be a

modal time step analysis and that it include the foIIowlng features:

• Three-dimenslonal modes obtained from finite element methods.

• Curved beam description of modes.

Large displacement equations of motion, to Include four or five

bending and twisting degrees of freedom with the capability of inc1udlng chordwise bending for future growth.

Complete induced flow capab111ty such that various methods of Inductlon can be selected, from momentum methods to vortex and pressure

potential methods.

Non-llnear aerodynamics for steady state operation, Includlng high

angles of attack.

Non-steady aerodyn_mlc effects to include non-steady coefficlents,

accounting for phaslng, to be added to the steady state descrlptlon

with the ability to substitute emplrlcal data or theory (syntheslzlng of data).

Three dimensional treatment of alrloads, including radial and interblade effects.

21
5_ It is also recommended that wind tunnel tests be conducted on twodimensional Prop-Fan airfoil sections, to provide data for use in

Improving the theoretlcal analyses.

tigatlons of the Followlng:

These tests should include InvesSteady state data.

Unsteady data (synthesis).

High angle of attack.

High Mach number effects (compressibility).

Effects of sweep,

22
.

.

.

REFERENCES

Goepner, B.W., "Static Flutter Tests of HSD Prop-Fan Models", United

Technologles Research Center Report R81-3354]4, February, 1982 (controlled circulation).

Turnberg, J., "Classical Flutter Stabillty of Swept Propellers." Proceedings of the AIAA/ASME/ASCE/AHS 24th Structures, Structura] Dynamics

and Materials Conference, May, 1983, Lake Tahoe, Nevada.

Stelnman, D.B., "Aerodynamic Theory of Bridge Oscillations", American

Society of Civil Engineers, Transactions, Paper No. 2420, October, 1949.

Smith, A.F., "Analysis and Test Evaluation of the Dynamic Response and

Stability Of Three Advanced Turboprop Models at Low Forward Speed", NASA

CR 175026, December 1985.

23/24

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25
TABLE I1: STRAIN GAGE DESIGNATION MODEL PROP-FAN

UTRC STATIC STALL FLUTTER TESTS

PROP-FAN RADIAL BLADE NO GAGE DESIGNATION

MODEL DESCRIPTION STATIONrlN. 1 2 3 _ 5

SR-2 Mid-Blade Bending 7.0 BG]-2 BG2-2 BGS-2 BG6-2

SR-2 Shear-V Gage 7.5 BGI-3 BG2-3 BGS-3 BG6-3

SR-2 Tip Bending lO.O BG1-4 BGS-4

6 7 8 _. lo

SR-3 Inbd. Bending 4.4 BGI-1 BGS-1

SR-_ Shear-V Gage 9.6 BGI-4 BG2-4 BGS-_ BG6-4

SR-3 Tlp Bending 10.7 BGI-6 BG2-6 BGS-6 BG6-6

SR-5 Inbd. Bending 5.3 BGI-I BG2-1 BG3-1 BG6-! BG7-1 BGB-1

SR-5 Shear-V Gage 8.9 BGI-5 BG6-5

SR-5 Tip Bending 10.4 BG1-3 BG6-3

26
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DESIGNATION

MSC NASTRAN

BESTRAN

H025

H027

H444

MODES

F214

F203

PLT203

TABLE IV

HAMILTON STANDARD COMPUTER PROGRAMS USED IN THE

STALL FLUTTER ANALYSIS CODE

PURPOSE

Finite element analysls used to predict

vibratory mode shapes and frequencies

for swept, thin structures.

Hamilton Standard finite element analysls used to predict vibratory mode

shapes and frequencies for swept, thin

structures.

Beam type analysis used to predict vlbratory bending mode shapes and frequencles for straight propeller blades.

Beam type analysis used to predict vibratory torsion mode shapes and frequencies for stralght propeller blades.

General Goldstein-_ype performance

strip analysis for propellers. Provides power, thrust, section force data

and angles of attacK. Section lift and

moment curve slopes are determined for

use in the stall flutter analysls, F203.

Converts mode shapes from finite element methods or beam methods to a beam

type description for use In the F203

flutter analysis.

Thls program transposes all co-ordlnate

system motions into the blade section

co-ordlnate system In order to take advantage of small angle assumptions.

Eigen-solution modal stability analysls. Calculates damping and frequency

using unsteady aerodynamics.

Plots the damping and frequency results

obtained in F203.

28
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in

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/

AIR E:XCHANGER

VALVES

/- •

f-/-- _/

PTR LOCATION I

A,N O R,V [_

Z,44M

_TEST sECTION

CIRCUT

5.79M _ o--l'O C K E D

÷ -_- ------

4.27M

t

5.49M DIFFUSER I

N

FIGURE 7 WIND TUNNEL CIRCUIT

36
_T\

A

r.

Z

u,J

Zo

u

la, wwoZa

...I

G. oLW

39
STATION 0'

N

STATtC

PRq_BE

SUPPORT

STRUTS ......_

TUNNEL/ I

SPANNING

RAKE

TOP VIEW

STATION

27g

STATION

289

ii

STATION 'O"

SIDE VIEW

TUNNEL

RAKE STATION

289

FIGURE 9 PROPELLER DYNAMOMETER INSTALLATION

38
ORIGINAL F._G_ tS

OF POOR OUALI'P(

i

f

_TUNNEL

SPANNING RAKE

II

STATIC PRESSURE

PROBE

FIGURE 10. PROP- FAN MODEL INSTALLATION LOOKING DOWNSTREAM

39
ORIGINAL PAC:_ IS

OF. POOR QUALITY

"N

STATIC

PROBE

.......

\

.=

FIGURE 11. PROP - FAN MODEL INSTALLATION LOOKING UPSTREAM

40
_L_T?_;'j

VIBRATORY

R £SPONS£ ANALYSES

BEAM I$R-2)

NASTRAN (SR-3&5)

I

FREQUENCIES

MODE SHAPES

MODAL MASSES

k F214CLIST

I

COORDINATE I

TRANSFORMATION I

I AERODYNAMIC

DATA BANK

|mu_m,,Imq _

.I STEADY STATE DATA

L:EXTENDEDINTO STALL1__,

H444 I

I PERFORMANCE I

F203CL CLIST

I i

_mmm_ mlmmm

I I

_mmmmmBJ l" TRANSFORMED I

I FREQUENCIES I

MODE SHAPES I

MODAL MASSES :

F203

STABILITY ANALYSIS

STALL FLUTTER

EIGENVALUE_ ENERGY

SOLUTION ;BALANCE

SOLUTION

PLT203

PLOT PROGRAM

_m m m mBm

DAMPING & FREQUENCY

VS.

BLADE ANGLE

I ...... I

FIGURE 12. BLOCK DIAGRAM FOR STALL FLUTTER ANALYSIS

:LIFT

ICURVE SLOPES

IAT OPERATING

CONDITIONS

41
DAMPING TO

CRITICAL

DAMPING RATIO

,02 ---

0.0

--.0Z D

THIRD MODE

7000 RPM

STABLE

SR-3-'_-_.

_._.... . \ \ ...-_'.;

-- UNSTABLE

I i I I i I 1 i I I

I 0 20 30 40 50

j_.75, BLADE ANGLE (DEG,)

FIGURE 13. MODEL PROP-FAN UTRC WIND TUNNEL

STATIC TESTS STABILITY PREDICTIONS

42
DAMPING TO

CRITICAL

DAMPING RATIO

m

.0Z

0.0

B

-.02

STABLE

UNSTABLE

THIRD MODE

9000 RPM

#_. _._'_ _-_

i 1 I I I I I l J

I0 2O 30 4O 5O

_.75, BLADE ANGLE {DEG.)

FIGURE 14. MODEL PROP-FAN UTRC WIND TUNNEL

STATIC TESTS STABILITY PREDICTIONS

43
DAMPING TO

CRITICAL

DAMPING RATIO

.02

0.0

!

--.02 B

STABLE

MODE: 7000 RPM

1ST

2ND

3RD_

4TH

UNSTABLE:

i i I i [ [ i [ [ I

I0 20 30 40 5u

_7S-BLADE: ANGLE-DE:G,

FIGURE 15. SR-3 MODEL PROP-FAN UTRC WIND TUNNEL

STATIC TESTS STABILITY PREDICTIONS
U IiJ 0 a Id

W

00 e_o

_oo

•ra0 i..

•, i.m +t,,,_ __N _- ,0o

N

o)

OOOOOO

OOOOO

UZ

Ill

DN

0%W

b_

N

[1:

?

0.

W_ !- I-

¢_0 J<<

13_ _. ,J

I_I V) laJW b,

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_

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IIIIi o_,

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ooooO

UZ

DNW

I1: h

i

0:

m

w? Oo

O_ _

o

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U.iOoD

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<_

in i- I O I I

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:[ :["_ :E

(2m []

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0.M

ILl

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45
-, ];

a

Id_ 0..o

,',_ .=._ umr

0__a IDff/ =uPo _/ a ._ = ¢ o ,/[_]| ti

(_ 9[ .I Id_ ,*=r_• "- _ u ,.. = _ Jlll_l

¢el _,_Oo_L_ I_

0__ ti ® o= g Z l I I I A2 0 O 0 0

1"0'=1(3] q'IONV qav18 ',.-13U_I

Z 0: < I,-

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m Id r_ m _1|1

z < _ DIlII

: °°" NIII

,n '=

+, . _J, NII

I I I I

0 ¢_ 0 0

_ q N

('O3a) 3"IONV 3OV1B ".-13_

O

O

O

X Z-IDm N

IE

I ! I !

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QO

('O'310) ]IIONV 3OV'Ig '1.,-13_

I

==o:

O

O

II

X

_" _" _III { 7_"

i

I I I Jo

0 0 0 "_

J

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0

lZ

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0

i-

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0

[,J

LOm

b.

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J

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_ ._o=

_ o_"o <w _u_O

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.-, %%

® % \,z__"_'_l//III!1.=/-x

_ °°°_/

0001

I I I I

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1 I I I

0

IZ

0

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w

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48
4O0

SR-2

Z

I

ld

0

n.

o

IFIk

<

Z

(/!

3OO

2OO

IO0

¢

RPM

'_w-8S00

0 20 3O 4O 5O 60

Z

I

w

0

0

P

IZ

4OO

30O

ZOO q

100 -_

o

IO

SR-3

RPM

75oo_- /

¢,-

2'

/

20 30 40 50 60

FIGURE 20.

400

_E

_, 300

I

bl

0

i¢ 200

0

P

Itl.

< 100

z

SR-5

I

010 20 30 40 50 60

_REF- REFERENCE BLADE ANGLE- DEG

TORQUE VS. REFERENCE BLADE ANGLE FOR THE UTRC PROP-FAN

STATIC TESTS, SR-2, SR-3 AND SR-5 MODELS

49
AVG.

VIBRATORY

ST R ESS

-+ kPa

30000

Z0000

10000

REF. BLADE ANGLE I' 36,2 °

THIRD MODE

5P

7000 RPM

REF. BLADE ANGLE " 36.2

TIPGAGE BGS-4

FIRST A

I /! MODE 2P 3P II I

0.00 0,10 0,20 0.30 0.40 0.50 0,60 0.70 0.80 0,90 1,00

FREQUENCY- kHz

FIGURE 21, $R-2 PROP-FAN MODEL BLADE STALL FLUTTER TESTS AT UTRC

5O
15000

10000

AVG.

VIBRATORY

ST R ESS

+- kPa

5000

THIRD MODE

oi

0.00 0.10

ZP

O.2O 0.30

8500 RPM

REF. BLADE ANGLE =31.5 °

SHEAR GAGE BGS-3

5P 6P

0,40 0.50 0,60 0.70 0,80 0.90 1.00

FREQUENCY- kHz

FIGURE 22. SR-2 PROP-FAN MODEL BLADE STALL FLUTTER TESTS AT UTRC

51
30000

Z0000

AVG.

VIBRATORY

STRESS

-+ kPa

10000

1ST MODE

ZP

5000 RPM

REF. BLADE ANGLE ,, 50.3 °

MID BLADE BENDING BGI-2

7P

5P

4P 6P -._ -J __ J

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00

FREQUENCY- kHz

FIGURE 23. SR-2 PROP-FAN MODEL BLADE STALL FLUTTER TESTS AT UTRC

.52
AVG.

VIBRATORY

ST R ESS

-+ kPa

15OOO

10000

5000

0

0.00

3P

9050 RPM

REF, BLADE ANGLE : 3 !.7 °

TIP BENDING GAGE BGI-6

SECOND

FIRST rl _ sp f I

0.10 0.20 0.30 0.40 0.50 0.60 0,70 0.80 0.90 1.00

FREQUENCY- kHz

FIGURE 24. SR-3 PROP-FAN MODEL BLADE STALL FLUTTER TESTS AT UTRC

53
18000

I0000

AVG.

VIBRATORY

STRESS

-+kPll

S000

0

0.00

FIGURE 25.

2ND

MODE

70 20 R PM

REF. BLADE ANGLE " 32.7"

TIP BENDING GAGE BGI-6

tP 2P 3P

I

0.10 0.20 0.30 0.40 0.50 0.80 0.70 0.80 0.90

FREQUENCY - kHz

SR-3 PROP-FAN MODEL BLADE STALL FLUTTER TESTS AT UTRC

3RD 5TH

MODE MODE

_4P 5P 4TH 1_

1.00

54
30000,_

ZOO00

AVG.

VIBRATORY

STRESS

+ kPa

!0000

|P

0.00

FIRST

MODE

ZP

5010 RPM

REF, BLADE ANGLE = 50.3 °

INBOARD BENDING GAGE BGI-I

0. l 0 0,20 0.30 0,40 0.50 0.60 0.70

FREQUENCY- kHz

- _L I

0.80 0.90 !.00

FIGURE 26. SR-3 PROP-FAN MODEL BLADE STALL FLUTTER TESTS AT UTRC

55
6000

4000

AVG,

VIBRATORY

ST R [5S

± kPa

2000

8";00 RPM

REF. BLADE ANGLE = 35.7 °

SHEAR GAGE BGI-3

2ND

MODE

IP 2P

! ST

MODE / _ /

0.00 0,10 0.20 0.30 0.40 0,50

3RD

MODE 4TH 5TH

MODE MODE

,I 6TH

0,60 0.70 0.80 0.90 I.O0

FREQUENCY- kHz

FIGURE 27. SR-5 PROP-FAN MODEL BLADE STALL FLUTTER TESTS AT UTRC

56
8000--

40OO

AVG.

VIBRATORY

ST R ESS

-+ kPa

2000

0

0.00

2ND

MODE

IST

6500 RPM

REF. BLADE ANGLE ,' 49.8 °

INBOARD BENDING GAGE BG2-1

MODE

3P 4P ' 5P , _P

0.10 0.20 0.30 0.40 0.50 0.80 0.70 0.80 0.g0 !.00

FREQUENCY- kHz

FIGURE 28. $R-5 PROP-FAN MODEL BLADE STALL FLUTTER TESTS AT UTRC

57
)REP

BLADE ANGLE

DEG.

80m

60m

40 D

20 --

0

0

1ST EXPERIMENTAL MODE

BGI-I 160-260 NZ

_"_44000

'_ _3000 INBOARD GAGE

\ ! 2000 STRESS (-+PSI)

IOOO I

| _.. 2000 \_'-_,ooo

/,ooo\_,ooo(_,;oo:

_ _ _\\_.OOOo

) _ ] I/_/ "-",,,Jz zooo

- o ,OOO

/ "zooo I

"/ 7500 TEST

OPERATING

LIMIT

L .__1 i I I

2000 4000 6000 8000 12000

RPM

REF

BLADE ANGLE

DEG.

B0

60 M

40_

Z0 --

2ND EXPERIMENTAL MODE

BGI-6 380.---_.-450 HZ

___ "qr_" 2ooo I o o5oo3O

2ooo

oOO

TIP GAGE

STRESS (-+ PSi)

-- ! 000

o i I I I I

0 2000 4000 6000 8000 ! 2000

RPM

FIG. 29 MODAL VIBRATORY STRESS (-+PSI)*

CONTOURS FROM THE SR-3 MODEL

BLADE STALL FLUTTER TESTS AT UTRC

*(6.89 X PSI =kPa)

58
REF

BLADE ANGLE

DEG

8O

6O

4O

Z0

0

0

3RD EXPERIMENTAL MODE

BG|-6 605-',-640 HZ

TIP GAGE STRESS (--- PSI)

1500

!000

500-- _'---- 1000

PREDICTIONS

I ! I I I

2000 4000 6000 8000 12000

RPM

_REF

BLADE ANGLE

DEG

8O

60 --

40 D

20

0

0

4TH EXPERIMENTAL MODE

BGI-4 670-755 HZ

SHEAR GAGE STRESS (+ PSI)

1000

CISO0

% .

1000

750

I I I I I

2000 4000 6000 8000 12000

RPM

FIGURE 29. (CONT°D)

59
REF

BLADE ANGLE

DEG.

80

60 m

40 D

20

o I

0 2000

5TH EXPERIMENTAL MODE

BGI-8 815 TO g00 HZ

,ooo/_._-._°°o°

°

1500

TiP GAGE

STRESS (+ PSI)

C 1000

I I ], I

4000 6000 8000 12000

RPM

FIGURE 29. (CONT'D)

6O
TOTAL

VIBRATORY

STRESS

± kPa

TOTAL

VIBRATORY

ST R ESS

± kPa

150000

100000

50000 D

0 o

IS0000

I00000

S0000

0

0

REFERENCE

BLADE ANGLE

31.8 NASA-LERC 10X 10

39.5 NASA-LERC 10 X 10

+_-F 32.0 STATIC TESTS (UTRC)

X_x 39.8 STATIC TESTS (UTRC)

BGI-ZMID*BLADE BENDING

x_XX / NASA

xJTRC

I 1_ I I I I I I ,I I

1000 2000 3000 4000 S000 6000 7000 6000 g000 10000

ROTATIONAL SPEED, RPM

REFERENCE

BLADE ANGLE BGI-3 SHEAR

31.8 NASA-LERC 10 X 10

39,5 NASA-LERC 10 X 10

+_+ 32.0 STATIC TESTS (UTRC}

X--x 39.8 STATIC TESTS (UTRC}

NASA

NASA

I _5_ ! I Z I I I I

1000 2000 3000 4000 S000 6000 7000 8000 9000 10000

ROTATIONAL SPEED, RPM

FIGURE 30. SR-2 MODEL PROP-FAN COMPARISON OF TESTS MACH NO, = 0.0 NO TILT

61
TOTAL

VIBRATORY

ST R ESS

±kPa

200000

150000

100000 --

50000 --

0

0

!

1000

BGI-4 TIP BENDING

X

NA_ UTRC

X

NASA .at

L_ i I .. _f BLADE ANGLE

x/._._" _ 39.5 NASA LEWIS, l0 X 10

/_ +/+ 31.5 STATIC TESTS (UTRC)

' ×_ 39.8 STATIC TESTS (UTRC)

I I I I L I I I I

2000 3000 4000 5000 6000 7000 8000' 9000 10000

ROTATIONAL SPEED, RPM

FIGURE 30._(CONT'D). SR-2 MODEL PROP-FAN COMPARISON OF TESTS

MACH NO. = 0.0 NO TILl"

62
SO000

TOTAL

VIBRATORY

ST R ESS

_+ kP}

UTRC

BG 14 SHEAR

NASA

UTRC

NASA

0

0 I000 ZOO0 3000

I

4000 $000 6000 7000

ROTATIONAL SPEED, RPM

I I I

8000 9000 10000

100000

TOTAL

VIBRATORY

STRESS

2kPa

30000

REFERENCE:

BLADE ANGLE

32.0 NASA LEWIS

O---"-O eo.o NSAA LEWIS

,'. ._ 31.7 UTRC

T T eo.o UZRC

BGI-8 TIP BENDING

o I I

0 1000 10000

UTRC

NASA

I I I I I I I I

2000 3000 4000 5000 6000 7000 8000 9000

ROTATIONAL SPEED, RPM

FIGURE 31. SR-3 MODEL PROP-FAN COMPARISON OF TESTS MACH NO. = 0.0 NO TILT

63
150000

100000

TOTAL.

VIBRATORY

STRESS

÷ kPa

50000 --

0

0

FIGURE 31.

m

REFERENCE

BLADE ANGLE

32.0 LEWIS

....... 50.0 LEWIS

+ I 31.I UTRC

X_XS0,5 UTRC

MACH I 0.0 NO TILT

UTRC

BGI-I INBOARD

BENDING

X

i l l I I I I I I I

1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

ROTATIONAL SPEED, RPM

CONT'D. SR-3 MODEL PROP-FAN COMPARISON OF TESTS MACH PJO. _: 0.0

NO TILT

6,t
TOTAL

VIBRATORY

STR ESS

± kPa

TOTAL

VIBRATORY

STRESS

± kPa

TOTAL

VIBRATORY

STR ESS

± kPa

100,000

S0,000

100,000

50,000

MACH NO. - 0,0 NO TILT

BGI-I INBOARD BENDING

S

UTRC

0 2000 4000 6000 8000 ! 0,000

ROTATIONAL SPEED, RPM

-- 32.0 NASA LEWIS _ _ (._UTRC NASA

---- $0.0 NASA LEWIS I _" t

['1 31.9 STATIC TESTS (UTRC) I - /

0 4,.6 STATIC TESTS(UTRC} .6 /

-- ,_)._f .,.,,,v _ "/

J,.-"

(__ -""'NASA

o i I I]]'a-FI I I I I I I I

0 2000 4000 6000 8000 10,000

ROTATIONAL SPEED, RPM

I00,000

50,000 m

0

0

BGI-5 CHORDWISE BENDING

/NASA

 TRCy

f_

UTRC

NASA"

_'_l I l I l i i i

Z000 4000 6000 8000 I 0,000

ROTATIONAL SPEED, RPM

FIGURE 32. SR-5 MODEL PROP-FAN COMPARISON OF TESTS

65/'66

APPENDIX A

TOTAL VIBRATORY STRESS" PLOTTED AS A FUNCTION

OF RPM FOR VARIOUS BLADE ANGLES AS OBSERVED IN

THE UTRC STATIC STALL FLUTTER TESTS ON THE SR-2,

SR-3 AND SR-5 MODEL BLADES

*rnfrequently repeating peak stress as taken from brush charts.

67/68

FIG NO

A-I

A-2

A-3

A-4

A-5

A-6

A-7

A-8

A-9

A-IO

A-II

A-12

A-13

A-14

A-15

A-16

A-17

A-I8

PROP-FAN

MODEL

SR-2

SR-2

SR-2

SR-2

SR-2

SR-2

SR-2

SR-2

SR-2

SR-3

SR-3

SR-3

SR-3

SR-3

SR-5

SR-5

SR-5

SR-5

APPENDIX A

TABLE OF CONTENTS

BLADE ANGLE,

DEG

15.8, 19.6

23.8, 27.4

29.7, 32

31.5, 36.2

39.8, 50.3

60.3, 69.9

79.6

15.8 through 32

36.2 through 79.6

]0.0 through 32.7

34.0 through 44.9

50.3 through 80.0

10.0 through 32.7

34.0 through 80.0

lO.O, 12.0

28.0, 39.1

35.7, 39.9

49.8, 59.6

STRESS TYPE

SHEAR

SHEAR

SHEAR

SHEAR

SHEAR

SHEAR

SHEAR

TIP-BENDING

TIP-BENDING

TIP-BENDING

TIP-BENDING

TIP-BENDING

TORSION

TORSION

INBD. BENDING

TIP BENDING

SHEAR

INBD. BENDING

TIP BENDING

SHEAR

INBD. BENDING

TIP BENDING

SHEAR

INBD. BENDING

TIP BENDING

SHEAR

69
FIG NO

A-19

A-20

A-21

APPENDIX A

TABLE OF CONTENTS (Contlnued)

PROP-FAN

MODEL

BLADE ANGLE,

DEG

SR-5 69.6, 79.7

SR-5

SR-5

10.0 through 59.6

59.6 through 79.7

STRESS TYPE

INBD. BENDING

TIP BENDING

SHEAR

INDB. BENDING

INDB, BENDING

70
40000

TOTAL

VIBRATORY

STRESS

-+ kPa

ZOO00

MEASURED VIBRATORY SHEAR STRESS

(INFREQUENTLY REPEATING PEAK

STRESS TAKEN FROM BRUSH STRIP CHARTS}

REF ,m ! 5.8 °

SHEAR STRESS

300

-- 2.00

(_ BG6-3

I'_ BGS-3

.m..m TORQUE .

m

1 100

o I I "---t "" T" 0

0 2000 4000 6000 8000 10000

RPM

TORQUE

FT-LBS

_REF " 19'6°

SHEAR STRESS "-'1 300

TOTAL 40000 D

VIBRATORY

ST R £SS

+- kPa O BGS-3 200

ll TORQUE

zoooo- 1

0_ 0

0 2000 4000 6000 8000 10000

RPM

FIGURE A-1. SR-2 8 WAY STATIC PROP-FAN TESTS AT UTRC

TORQUE

FT-LBS

71
40000

TOTAL

VIBRATORY

ST R ESS

± kPa

Z0000_

MEASURED VIBRATORY SHEAR STRESS

{INFREQUENTLY REPEATING PEAK

STRESS TAKEN FROM BRUSH STRIP CHARTS)

REF" m Z3.8

SHEAR STRESS

_ TORQUE

C) BGS-3

/

/

J

f

S

o i

0 2000 4000 6000" 8000

RPM

300

-- 2OO

m

I00

I0000

TORQUE

F.T-LB

40000

TOTAL

VIBRATORY

ST R ESS

+ kPa

20000

REF" - 27.4

SHEAR STRESS

_ TORQUE

C) BGS-3

/

/f

IS I i I I i l i

2000 4000 6000 8000

RPM

-- 300

200

m

I00

10o0o

Q

TORQUE

F.T-LB

FIGURE A-2. SR-2 8 WAY STATIC PROP-FAN TESTS AT UTRC

72
40000

TOTAL

VIBRATORY

ST R ESS

+k Pe

20000

n i

0

MEASURED VIBRATORY SHEAR STRESS

INFREQUENTLY REPEATING PEAK

STRESS TAKEN FROM BRUSH STRIP CHARTS}

SHEAR STRESS

/

/

O BGT-3

_] BGS-3

/

J

/

/

I I I I I I I I

2000 4000 6000 8000

RPM

-- 300

ZOO

-- 100

0

10000

Q

TORQUE

FT-LB

40000

TOTAL

VIBRATORY

5TRESS

kPI

20000 B

0

0

SHEAR STRESS

_REF - 32 °

//

/ --- TORQUE

f _// 0 BGT*3

t / [] BG5-3

I I I l l.,. I ! I I

2000 4000 8000 8000

RPM

FIGURE A-3. SR-2 8 WAY STATIC PROP-FAN TESTS AT UTRC

3O0

Q

-- TORQUE

FT-LB

200

100

0

10000

73
MEASURED VIBRATORY SHEAR STRESS

(INFREQUENTLY REPEATING PEAK

STRESS TAKEN FROM BRUSH STRIP CHARTS)

SHEAR STRESS

40000

TOTAL

VIBRATORY

ST R ESS

-+ kPu

20000_

2000 4000

RPM

/

/

/

/

I I I_ I I I

6000 8000

m 3oo

m

200

Q

TORQUE

m

FT-LB

-- I00

0

I0000

40000

TOTAL

VIBRATORY

ST R ESS

+kPa

20000 / /

J/////.(:7"" _ I/

t 1_._ "_ 0I" I I I I I J

0 2000 4000 6000

RPM

FIGURE A-4. SR-2 8 WAY STATIC PROP-FAN TESTS AT UTRC

l

J I o

8000 !0000

300

20O

Q

TORQUE

FT-LB

100

74
40000

TOTAL

VIBRATORY

STRESS

--- kPa

20000 1

MEASURED VIBRATORY SHEAR STRESS

(INFREQUENTLY REPEATING PEAK

STRESS TAKEN FROM BRUSH STRIP CHARTS)

REF " 39.8

-_ TORQUE

O BG6-3

////

ooi//

f

11 I I I I I I !

2000 4000 6000 8000

RPM

300

ZOO

100

0

10000

Q

TORQUE

FT-LB

40000

TOTAL

VIBRATORY

STRESS

+ kPa

20000 m

0_

0

SHEAR STRESS

J_ REF " 50,3

------ TORQUE

O BG6-3

/

//

/

I I I I I I I I

2000 4000 8000 8000

RPM

300

200

100

0

10000

FIGURE A-5. SR-2 8 WAY STATIC PROP-FAN TESTS AT UTRC

Q

TORQUE

FT-LB

75
40000

TOTAL

VIBRATORY

STRESS

:1: kPa

20000

MEASURED VIBRATORY SHEAR STRESS

(INFREQUENTLY REPEATING PEAK

STRESS TAKEN FROM BRUSH STRIP CHARTS)

_REF ,B 60.3 SHEAR STRESS

I I I I 1 I [ I I

0 ZOO0 4000 6000 8000

RPM

- 300

Q

-- TORQUE

FT-LB

-- 200

-- I00

0

I0000

40000

TOTAL

VIBRATORY

STRESS

÷ kPa

20000 m

0

0

------TORQUE

_REF-69.9 SHEARSTRESS O BG6-3

/

I l t I J I i i l

2000 4000 6000 8000

RPM

FIGURE A-6. SR-2 8 WAY STATIC PROP-FAN TESTS AT UTRC

-- 300

Q

TORQUE

FT-LB

-- 200

n

-- I00

10000

?6
40000

TOTAL

VIBRATORY

ST R ES5

+ kPa

20000

m

o I

o

FIGURE A-7.

MEASURED VIBRATORY SHEAR STRESS

(INFREQUENTLY REPEATING PEAK

STRESS TAKEN FROM BRUSH STRIP CHARTS)

REF " 79.6" SHEAR STRESS

om TORQUE

O BG6-3

I f I I I I I

Z000 4000 6000 8000

RPM

SR-2 8 WAY STATIC PROP-FAN TESTS AT UTRC

-- 3OO

-- 200

m

-- 100

o

10000

Q

TORQUE

FT-LB

7?
TOTAL

VIBRATORY

STRESS

±kPa

I00000

0 o

MEASURED VIBRATORY STRESS (INFREQUENTLY

REPEATING PEAK STRESS TAKEN FROM BRUSH STRIP CHARTS)

REF BLADE TiP BENDING STRESS

ANGLE, DEC BG-$-4

@ Is.a"

50000

100000

0 19,6

23.8

z7.4

/_ z9.7

(_ 31.5

2000

TOTAL

VIBRATORY 50000

STRESS

-+ kPa

REF BLADE

ANGLE. DEG

m 0 31.5

O3z

Y

I

2000

°j

I I i I

4000 6000 8000 10000

RPM

TIP BENDING STRESS

BG 5-4

I I I I

0 0 4000 6000 6000 | 0000

RPM

FIGURE A-8. SR-2 MODEL PROP-FAN TEST AT UTRC

'78
TOTAL

VIBRATORY

STRESS

±kPa

290,000 m

7.00,000

150,000

! 00,000

50,000

MEASURED VIBRATORY STRESS

(INFREQUENTLY REPEATING PEAK STRESS

TAKEN FROM BRUSH STRIP CHARTS)

REF. BLADE

ANGLE, DEG

36,Z

39..

[] 50.3

0 60.3

69.9

_7 79.6

L [ I [,,, I I I I

0 2000 4000 6000 8000

RPM

FIGURE A-9. SR-2 MODEL PROP-FAN TESTS AT UTRC

TIP BLENDING STRESS

BG-S-4

I I

I 0,000

79
MIrABURED VIBRATORY STRESS

(INFREQUENTLY REPEATING PEAK STRESS

TAKEN FROM BRUSH STRIP CHARTS)

TOTAL

VIBRATORY

ST R ESS

±kPa

40000

20000

0

0

AVERAGE OF BG t-6, 5-6, 2-8 AND 6-6

TIP BENDING

BETA R EF-DEG

10.0

• 12.0

_mum 15.9

"" -- _ 19.9

1000 2000 3000 4000 5000 6000 7000 8000 9000

ROTATIONAL SPEED-RPM

100000

80000

TOTAL

VIBRATORY

STRESS

_+ kPa

60000

40000

20000

0

0

AVERAGE OF BG 1-6, 5-6, 2-6, AND 6-6

TiP BENDING

'='ETAREF-DEG ./'\ /

23.6 /

,6 /\. / /

3,.7 / \ / __. /

..... 32.7 /_ • -- "_, • *_

; / \/

!/

I

I000

._..JI ..- ....

1 I 1 I I I i I

2000 3000 4000 5000 8000 7000 8000 9000

ROTATIONAL SPEED- RPM

FIGURE A-10. SR-3 MODEL STATIC TESTS AT UTRC

80
140000

120000

100000 D

TOTAL

VIBRATORY

ST R ESS

+ kPa

80000

60000

40000

20000 ,---

0

0

MEASURED VIBRATORY STRESS

(INFREQUENTLY REPEATING PEAK STRESS

TAKEN FROM BRUSH STRIP CHARTS

AVERAGE OF BG 1-6, 5-6, 2-6, AND 6-6

TIP BENDING

BETA REF*DEG

34.0

35.7

40.0

44.9

I

I

I

I

I

I

I

, I

I

I

I

I

I

/

I

/

I

/

/

/

/

/

J

/ /

/

^ /

I I I I i I

1000 2000 3000 4000 5000 6000

ROTATIONAL SPEED - RPM

I

7000

FIGURE A-I 1. SR-3 MODEL STATIC TESTS AT UTRC

81
MEASURED VIBRATORY STRESS

(INFREQUENTLY REPEATING PEAK STRESS

TAKEN FROM BRUSH STRIP CHARTS)

120000

100000

80000

TOTAL

VIBRATORY

STRESS

+ kPa

60000

40000

20000

0

0

BETA REF-DEG

50.3

60.0

69.9

m,_,, .... 80.0

AVERAGE OF BG 1-6, 5-6, 2-6 AND 6-6

TIP BENDING

I I I I I i I

1000 2000 3000 4000 5000 6000 7000

ROTATIONAL SPEED - RPM

FIGURE A-I2. SR-3 MODEL STATIC TESTS AT UTRC

82
MEASURED VIBRATORY STRESS

(INFREQUENTLY REPEATING PEAK STRESS

TAKEN FROM BRUSH STRIP CHARTS)

TOTAL

VIBRATORY

ST R ESS

+kPa

40000

20000 --

0

0

AVERAGE OF" BG !-4, 5-4. 2-4 AND 6-4

SHEAR GAGE- TORSION

BETA REF-DEG

10,0

" 12.0

u o .,,, mm 1"5,9

eme_m.m, ne-- 19.9

.... _.,,,,,.qr',-_ _ _,.,_m....,_ _ r'''ur_

I I_,qua___ --_i* -'-'-'-'-'-'-'-'-'!E _' -_az"-_ '..'2 'Zi_- _- .... "t ...... I......

1000 2000 3000 4000 SO00 $000 7000 8000 9000

ROTATIONAL SPEED-RPM

TOTAL

VIBRATORY

ST R ESS

+ kPa

40000 --

20000

0

0

AVERAGE OF BG I-4, 5-4, 2-4 AND 6-4

SHEAR GAGE: TORSION

BETA REF-DEG

" 23.6

27.6 o._ , .,,, ,'S° "''" ° "" " ""

i 31.7 ,._'_ .,,, .....- "" "-- -- --

IO00 2000 3000 4000 5000 6000 7000 8000 9000

ROTATIONAL SPEED-RPM

FIGURE A-! 3. SR-3 MODEL STATIC TESTS AT UTRC

83
MEASURED VIBRATORY STRESS

(INFREQUENTLY REPEATING PEAK STRESS

TAKEN FROM BRUSH STR|P CHARTS

TOTAL

VIBRATORY

STRESS

_kPa

60000 --

40000 --

20000 --

0

0

AVERAGE OF'BG 1-4, 5-4, 2-4 AND 6-4

SHEAR GAGE- TORSION

BETA REF*DEG

34.0

-. - -','--,.- 35.7

.,-- -,- -,- m 40.0

--- ,m,_, 44.9

I I I I I [ i

! 000 Z000 3000 4000 5000 6000 7000

ROTATIONAL SPEED-RPM

TOTAL

VIBRATORY

STRESS

±kPa

40000 u

20000 m

0

AVERAGE OF BG I-4. 5-4. 2-4 AND 8-4

SHEAR GAGE- TORSION

BETA REF-DEG

" 50.3

'" ...... 60.0

.,,D m _ _ 89.9

"" *m*-'- 80.0

e_ e J_,

I ! J

1000 2000 3000 4000 5000 6000

ROTATIONALSPEED-RPM

FIGURE A-14. SR-3 MODEL STATIC TESTS AT UTRC

I

7000

84
MEASURED VIBRATORY STRESS

(INFREQUENTLY REPEATING PEAK STRESS

TAKEN FROM BRUSH STRIP CHARTS)

40,000

TOTAL

VIBRATORY

ST R E55

-+ kPa

20,000

0

0

_REF " fO°

O SGI-3 TIP BENDING

r_ BGI-I INBOARD BENDING

_ BG1-5 SHEAR

2000 4000

RPM

I I I I

6000 8000 10,000

40,000

TOTAL

VIBRATORY

STRESS

±kPa

20,000

0

0

F

J3RE F : I 2 °

(_ BGI-3 TIP BENDING

C] BGI-I INBOARD BENDING

BGI-5 SHEAR

i I 1

2000 4000 6000 8000

RPM

I

10,000

FIGURE A-!5. SR-5 10 WAY STATIC PROP-FAN TESTS AT UTRC

8.5
40,000

TOTAL

VIBRATORY

STR ESS

+ kPa

20,000

0

0

40,000 m

TOTAL

VIBRATORY

STR EBB

+ kPa

20,000 --

0

0

MEASURED VIBRATORY STRESS

(INFREQUENTLY REPEATING PEAK STRESS

TAKEN FROM BRUSH STRIP CHARTS)

_REF " 28e

(_ BGI-3 TIP BENDING

[_ BGI-I INBOARD BENDING

O BGI-5 SHEAR

I

2000 4000 6000 8000

RPM

J

10,O00

REF " 31.9"

0 BGI-3 TIP BENDING

_J BGI-; INBOARD BENDING

_ BGI-5 SHEAR

S

! 1 1 I I I 1 I I

2000 4000 6000 8000

RPM

I

10,000

FIGURE A-16. SR-5 10 WAY STATIC PROP-FAN TESTS AT UTRC

86
UNITED

TECHNOLOGIES

60000 --

w

40000 m

VIBRATORY

ST R ESS

-+kPa

20000

0

m

60000m

40000m

VIBRATORY

STRESS

KPa

2000©_

0 ,

MEASURED VIBRATORY STRESS

{INFREQUENTLY REPEATING PEAK

STRESS TAKEN FROM BRUSH STRIP CHARTSI

REF : 35'7°

_ _r [-_ BGI-t INBOARD BENDING

BGI-5 SHEAR

I I I I I I I I I I

2000 400C (SO00 0000 10,000

RPM

(_ BG1-3 TIP BENDING

I'7 BGI-I INBOARD BENDING

BGI-5 SHEAR

I 1 J I I 1 I I I I

2000 4000 6000 8000 I 0,000

RPM

FIGUREA-17. SR-5 IOWAYSTATICPROP-FANTESTSATUTRC,

87
!00000

80000

60000

VIBRATORY

STRESS

± kP8

40000

20000

60000

40000

VIBRATORY

STRESS

± kPa

20000

MEASURED VIBRATORY STRESS

(INFREQUENTLY REPEATING PEAK

STRESS TAKEN FROM BRUSH STRIP CHARTS)

(_ BG1-3 TIP BENDING

[3 BGI-I INBOARD BENDING

_ BGI-5 SHEAR

I I l I I I

2000 4000 6000

RPM

¢

- /

I I I I

8000 I 0,000

O BG1-3 TIP BENDING

E] BGI-I INBOARD BENDING

BGI-5 SHEAR

t ! I l I I I I I I

2000 4000 6000 8000 I 0,000

RPM

FIGURE A-18. SR-5 10 WAY STATIC PROP-FAN TESTS AT UTRC

r

88
60000

40000

VIBRATORY

ST R ESS

+ kPa

Z0000

0

40000

VIBRATORY

ST R ESS

-+ kPa

20000

MEASURED VIBRATORY STRESS

(INFREQUENTLY REPEATING PEAK

STRESS TAKEN FROM BRUSH STRIP CHARTS)

C) BG1-3 TiP BENDING

[] BGI-! INBOARD BENDING

BGI-5 SHEAR

I I L I I I

Z000 4000 6000

RPM

I I I

8000

REF = 79"7°

/

I l I 1

2000 4000

O BGI-3, TIP BENDING

r7 BGI-I INBOARD BENDING

BGI-5 SHEAR

I

I0.000

1 I I I I I

6000 8000 I0,000

RPM

FIGURE A-19. SR-5 10 WAY STATIC PROP-FAN TESTS AT UTRC

89
100000

800O0

60000

TOTAL

VIBRATORY

ST R ESS

+ kPa

4000n

20000

m _ RE:F, DEG

[_ I0.0

28.0

[_ 31.9

35.7

O 39.9

C_3 49.8

59.6

2O0O

MEASURED VIBRATORY STRESS

(INFREQUENTLY REPEATING PEAK STRESS

TAKEN FROM BRUSH STRIP CHARTS)

GAGEBG2-1

INBOARD BENDING

b

FIGURE A-20. SR-5

! I 1

4000 8000 8000

J3 RPM

10 WAY STATIC PROP-FAN TESTS AT UTRC

]

10000

9O
MEASURED VIBRATORY STRi"..SS

(INFREQUENTLY REPEATING PEAK

STRESS TAKEN FROM BRUSH STRIP CHARTS)

I 2,000

I 0,00q

8000

±

VIBRATORY

STRESS

PSi

8000 m

m

400( B

ZOO0

0

0

GAGE BG2-1

INBOARD BENDING

_R EF

r-] 59.6

A e96

0 79.7

I I I I I I I ]

2000 4000 6000 8000

RPM

I

fO,O00

FIGUREA-21. SR-5 10 WAY STATIC PROP-FAN TESTS AT UTRC

91/92

APPENDIX B

STRESS PEAK TABULATION FOR THE SR-3 MODEL PROP-FAN

This table contalns data obtained from spectral analyses using the computerized peak picking rout|nes developed by Hamilton Standard. Listed are the

predominant frequencles measured for each straln gage signal, followed by the

stress amplitude. These are listed for each operating condition defined by"

REF. BLADE ANGLE

RPM

TORQUE

POWER COEFFICIENT

93/'94

ORdiNAL PAGE _S

POOR QUAUTY

Run

NO.

2

3

4

,a

6

7

8

9

10

11

12

13

14

15

16

2

3

4

5

6

7

8

9

10

11

12

13

]4

15

16

4

5

6

7

N

9

10

U

12

13

14

15

t6

18

19

20

21

22

-23

24

25

26

27

28

29

3O

31

32

18

19

20

TABLE B-I. SR-3 PROP-FAN MODEL STATIC TESTS AT UTRC

SPECTRAL STRESS PEAKS AND FREQUENCIES

Ref.

Blade

Angle Power

RP.._MM_ CoefF

12 9050 BO ,2617

12 8630 70 .2518

12 8025 65 ,2704

12 7500 55 .262

12 7025 48 .2606

12 6500 42 .2663

12 6025 35 .2583

12 5500 30 ,2657

12 5000 25 .2679

12 4490 20 .2658

12 3990 18 .3029

12 3525 12 ,25E7

12 3000 10 .2977

12 2550 9 .3708

12 2140 7 .4095

t2 9050 80 .2617

12 8630 70 ,2518

12 0025 65 .2704

12 7500 55 .262

12 7025 48 .2606

12 6500 42 .2663

12 6025 35 .2583

12 5500 30 ,2657

12 5000 25 .2679

12 4490 20 .2658

12 3990 18 .3029

12 3525 12 .2587

12 3000 10 .2977

12 2550 9 .3708

12 2140 7 .4095

12 9050 80 .2617

12 8630 70 ,2518

12 8025 65 ,2704

12 7500 55 .262

12 7025 48 ,2606

12 6500 42 ,2663

12 _025 35 ,25B3

12 5500 30 ,2657

12 5000 25 .2679

12 4490 20 ,2658

12 3990 18 ,3029

12 3525 12 ,25B7

12 3000 10 ,2977

12 2550 9 ,3708

12 2140 7 .4095

15.9 9030 115 ,3778

15.9 8540 105 .3857

15.9 8015 95 .3962

15.9 7540 80 ,377

15.9 7040 70 ,3784

15.9 6520 62 ,3907

15.9 6012 55 .4077

15.9 5520 45 ,3957

15.9 5030 40 .4236

15.9 4515 32 ,4206

15.9 4005 25 ,4176

15.9 3510 20 ,4349

15.9 3023 15 ,4398

15.9 2505 10 ,4269

15.9 2140 8 ,468

15.9 9030 115 ,3778

15.9 8540 105 .3857

15,9 8015 95 .3962

Gage No. of

No. Peak_ Spectral

881-4 0

881-4 0

881-4 0

881-4 0

BQI-4 O

8(]1-4 0

801-4 0

801-4 0

801-4 0

801-4 0

BGI-4 0

901-4

801-4

801-4 0

801-4 0

801-1 0

801-I 0

801-1 0

801-1 0

801-I 1 234 1128

801-1 I lOB 677

801-1 0

861-1 0

801-1 0

801-I 0

801-1 0

801-1 0

801°1 0

B01-1 0

BGI-1 0

B01-6 1 454 1011

Bn1-6 0

BG1-6 1 134 502

881-6 0

801-6 0

BG1-6 1 108 583

801-6 0

BG1-6 0

801 -_

801-6 0

801-6 0

801-6 0

B81-6 0

901-6 o

BGi -_ o

801-4 o

B01-4 0

801-4 0

B01-4 0

801-4 o

BG1-4 0

801-4 0

B81-4 0

B01-4 0

801-4 0

BG1-4 o

901-4 0

801-4 0

_01-4 0

BQI-4 o

801-4 0

801-4 0

B01-4 0

Frequencles(HZ)/Vlbratorv Stress (osl)

95
ORIGINALPAGE IS

OF POORQUALITY

Run

No.

18

19

20

21

22

23

24

25

26

27

29

29

30

31

32

18

19

20

21

22

23

24

25

26

27

29

29

30

31

32

33

34

35

36

37

39

39

40

41

42

43

44

45

46

47

33

34

35

36

37

-38

39

40

41

42

43

44

33

34

35

36

37

-38

39

40

41

TABLE B-1 (CONTINUED)

ReF.

Blade

Angle Power Gage No. of

Deg. RP._H Tor_._.rrrr_ Coeff No. Peaks

15.9 9030 115 .3778 B01-1 1

15,9 8540 105 ,3857 B81-1 0

15.9 8015 95 .3962 BGI-I 0

[5.9 7540 80 .377 961-1 2

15.9 7040 70 ,3784 DO1-[ 2

15.9 6520 62 .3907 981-I 2

[S,9 6012 55 .4077 B01-1 0

15.9 5520 45 ,3957 B01-1 0

15.9 5030 40 ,4236 _61-1 0

15,9 4515 32 ,4206 B81-1 0

15.9 4005 25 .4176 981-1 0

15.9 3510 20 .4349 BGI-1 0

15.9 3023 15 .4398 BGI-1 0

15.9 2505 10 .4269 B81-1 0

15.9 2140 8 .468 901-[ 0

15.9 9030 115 .3778 981-6 2

15.9 9540 105 .3857 981-6 I

15.9 B015 95 .3962 901-6 I

15.9 7540 80 .377 981-6 I

15.9 7040 70 .3794 BG1-6 |

15.9 6520 62 .3907 BGI6 2

15.9 6012 55 ,4077 BGI-6 0

15,9 5520 45 .3957 981-6 o

15.9 5030 40 .4236 BGI-6 0

15.9 4515 32 .4206 901-6 o

15.9 4005 25 .4176 981-6 0

15.9 3510 20 .4349 901-6 0

15.9 3023 15 ,4398 981-6 0

15.9 2505 10 ,4269 901-6 0

15o9 2140 8 ,469 B01-6 0

19.9 9020 165 .5433 981-4 0

[9.9 8500 150 ,5562 901-4 0

19.9 8000 135 .5651 B81-4 0

[9.9 754_ 120 ,5652 981-4 0

19,9 7030 100 .542[ 901-4 0

19.9 6525 92 .5789 BGI-4 0

19.9 6010 75 ,5563 981-4 0

19.9 5520 60 ,5276 B81-4 0

19.9 5025 50 ,5305 BG1-4 0

i9,9 4510 42 ".5532 901-4 0

19,9 4010 35 .5831 B81-4 0

19,9 3510 29 ,6306 BGI-4 0

19,9 3000 21 .6251 981-4 0

19.9 2500 15 .643 981-4 0

19.9 2140 11 .6435 B81-4 0

19.9 9020 165 ,5433 861-1 1

19.9 8500 150 .5562 881-1 2

19.9 8000 135 .5651 861-I I

19.9 7542 120 ,5652 861-1 0

19.9 7030 I00 .5421 881-1 2

19.9 6525 92 ,5789 901-1 2

19.9 6010 75 ,5563 881-1 1

19.9 5520 60 ,5276 BOJ-I 0

19.9 5025 50 .5305 981-1 o

19.9 4510 42 ,5532 BGl-I 0

19.9 4010 35 .5831 901-I 0

19.9 3510 29 .6306 801-1 0

19.9 9020 165 .5433 9G1-6 2

19.9 8500 150 .5562 981-6 1

19.9 8000 135 ,5651 901-6 0

19.9 7542 120 .5652 801-6 0

19.9 7030 I00 .5421 88[-6 [

19.9 6525 92 .5789 88[-6 1

19.9 6010 75 .5563 80[-6 o

19.9 5520 60 .5276 88[-6 0

19.9 5025 50 .5305 BOi-6 0

Spectral Frequencles(HZ)lVlbratory, Stress (ps;)

148 521

126 558 252 568

118 812 234 [243

108 786 218 982

148 749 448 692

142 520

134 560

126 619

118 894

108 690 218 588

452 659

142 552 284 521

266 529

118 657 234 862

108 759 218 688

200 509

150 582 452 860

142 532

118 600

108 678

96
ORIGINAL PAGE IS

OF POOR QUALITY

Run

No.

42

43

44

53

54

55

56

57

58

59

60

61

62

63

53

54

55

56

57 "

58

59

60

61

62

63

64

65

66

67

53

54

55

56

57

58

59

6O

61

62

63

64

53

54

55

56

57

58

59

60

61

62

63

64

65

66

67

53

56

55

56

57

58

59

60

61

62

Ref.

Blade

Angle

19.9 4510

19,9 4010

19.9 3510

23.6 9035

23.6 8570

23.6 8025

23.6 7520

23.6 7010

23.6 6540

23,6 5970

23.6 5535

23.6 5000

23.6 450O

23,6 4025

23,6 9035

23.6 8570

23.6 8025

23.6 7520

23.6 7010

"23.6 6540

23.6 5970

23.6 5535

23,6 5000

23.6 4500

23.6 4025

23.6 3510

23.6 2995

23.6 2510

23.6 2080

23,6 9035

23.6 8570

23.6 8025

23.6 7520

23.6 7010

23,6 6540

23.6 5970

23.6 5535

23.6 5000

23.6 4500

23.6 4025

23.6 3510

23.6 9035

23.6 R570

23.6 8025

23.6 7520

_3°6 7010

23.6 6540

23.6 5970

23.6 5535

23.6 5000

23.6 4500

23.6 4025

23,6 3510

23.6 2995

23.6 2510

23.6 2080

23.6 9035

23.6 8570

23.6 8025

23.6 7520

23.6 7010

23.6 6540

23.6 5970

23,6 5535

23.6 5000

23.6 4500

TABLE B-1 (CONTINUED)

Power Gage Ho. of

Torque Coeff No. Peaks Spectral Frequencles(HZ)/Vlbratory Stress (psi)

42 .5532 8191-6 0

35 .5831 801-6 0

29 .6306 801-6 0

230 .7549 BG1-4 0

205 .7478 901-4 0

180 °7488 801-4 0

155 .7343 801-4 0

130 .7088 861-4 0

120 .7517 801-4 0

100 .7517 801-4 0

85 °7433 801-4 0

70 .7502 801-4 0

55 .7277 801-4 0

45 .7442 BG1-4 0

230 .7549 801-1 1 452 553

205 .7478 861-I 1 142 542

180 .7488 801-1 0

155 .7343 801-I 0

130 .7088 BGI-I 2 I18 647 234 819

120 ,7517 B01-I 2 108 767 218 900

100 .7517 BGI-! I 200 551

85 .7433 BGJ-I 0

70 °7502 B01-1 0

55 .7277 801-I 0

45 .7442 801-1 0

35 .7611 B01-1 0

25 .7467 B01-1 0

18 .7654 B01-1 0

I0 .6192 861-1 0

230 .7549 801-6 2 150 527 452 983

205 .7478 BGI-6 I 142 580

180 ,7488 B01-6 0

155 ,7343 801-6 0

130 .7088 BGI-6 I 118 612

120 ,7517 801-6 I 108 707

100 ,7517 B01-6 0

85 ,7433 801-6 0

70 ,7502 801-6 0

55 ,7277 801-6 0

45 .7442 801-6 0

35 ,7611 BGI-6 0

230 .7549 B01-I 0

205 ,7478 B01-I 0

100 .748B 801-1 1 134 688

155 .7343 BGI-t I 126 033

t30 ,708B BGI-i I 118 1134

120 .7517 B01-1 2 110 792 21B 1469

100 .7517 801-1 2 100 659 200 520

B5 .7433 BGI-I I 92 540

70 .7502 B01-1 0

55 ,7277 801-1 0

45 .7442 801-I 0

35 ,7611 801-I 0

25 *7467 801-I 0

18 ,7654 801-1 0

10 .6192 901-1 0

230 ,7549 801-6 2 150 578 454 811

205 ,7478 801-6 I 144 682

180 .7488 BGI-6 1 134 860

155 .7343 801-6 1 126 970

130 .7088 B01-6 1 118 1031

120 .7517 801-6 1 II0 621

100 .7517 801-6 I I00 552

85 .7433 801-6 I 92 516

70 .7502 BGI-6 0

55 .7277 801-6 0

97
Run

No.

63

64

65

66

67

68

69

70

71

22

73

74

75

76

77

78

79

80

81

82

68

69

70

71

72

73

?4

75

76

77

78

79

80

81

82

68

69

70

71

72

?3

74

75

76

77

78

79

79

80

81

82

83

84

84

83

85

86

87

88

89

90

91

92

93

94

95

Ref.

Blade

Angle

23.6

23.6

23.6

23.6

23,6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27,6

27,6

27.6

27.6

27.6

27,6

27,6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27,6

27.6

27.6

_7.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

27.6

31.7

31.7

31.7

31,7

31.7

31.7

31.7

31.7

31.7

31.7

31.7

31.7

3L.7

31,7

31.7

RP___.

4025

3510

2995

2510

2080

9000

8555

8050

7585

6960

6570

6015

5520

4980

4530

4040

3505

3020

2570

2145

9000

8555

8050

7585

6960

6570

6015

5520

4980

4530

4040

3505

3020

2570

2145

9000

8555

8050

7585

6960

6570

6015

5520

4980

4530

4040

3505

3505

3020

2570

2145

9050

8570

8570

9050

8055

7535

7030

6505

6010

5505

5005

4520

4055

3550

3065

45

35

25

18

10

275

250

220

190

165

145

125

105

90

72

55

45

30

25

15

275

250

22O

190

165

145

125

105

90

72

55

45

30

25

15

275

250

220

190

165

145

125

105

90

72

55

45

45

30

25

15

290

260

260

290

240

210

180

155

135

115

95

75

65

45

35

Power

Coeff

.7442

.7611

.7467

.7654

,6192

.9096

.9151

.9095

.0848

.9126

.9195

.9256

.9232

,9722

,94

,9028

,9814

,8812

1.0141

,8734

,9096

.9151

.9095

.8848

.9126

,9195

.9256

,9232

.9722

,94

.9028

,9814

,8_12

1.0141

.8734

.9096

.9151

,9095

,8848

.9126

.9195

.9256

.9232

.9722

,94

.9028

.9814

.9814

.8812

1.0141

.0734

.9486

.9404

.9484

.9486

.991

.9909

.9758

,9814

1.0013

1.0167

1,016

.9835

1.0591

,9566

,9902

ORIGINAL PAGE IS

OF. POOR QUALITY

TABLE B-! CONTINUED

Gage

No.

901-6

801-6

901-6

801-6

901-6

881-4

BG1-4

901-4

801-4

801-4

901-4

801-4

901-4

B01-4

B01-4

801-4

B01-4

801-4

BG1-4

881-4

801-1

_01-1

881-1

BGI-1

B01-1

881-1

981-1

881-1

BGI-1

.BGI-I

881-I

881-1

801-I

BGI-1

881-1

801-6

BGI-6

881-6

BGI-6

801-6

BGI-6

BGI-&

881-6

881-6

BGI-6

BGI-6

881-6

B01-6

801~6

BOl-6

BOl-6

_01-4

981-4

B81-4

801-4

801-4

B81-4

801-4

801-4

901-4

901-4

801-4

901-4

801-4

881-4

BO1-4

No. of

Peak__s

0

0

o

o

o

o

o

0

0

0

0

0

o

0

0

o

o

o

o

o

o

2

2

1

g

0

1

I

o

0

o

o

0

o

2

2

SEectral

142 849 286 701

134 561 268 504

126 659

116 709 232 711

110 744 220 1644

184 596

166 633

150 890 452 1028

142 1096 428 957

134 778

134 661

116 751

110 518 220 618

740 680

752 737

734 738

722 643

716 625

704 736

694 760

686 612

680 571 684 554

Frequencles(HZ)/Vlbratory Stress (psi)

98
ORIGINAL PAGE |3

OF POOR QUALITY

Ref.

Blade

Run Angle

No_._.. Deg. RPH

96 31.7 2525

97 31.7 2130

--83 31,7 9050

84 31.7 8570

85 31.7 8055

86 31.7 7535

--57 31.7 7030

88 31.7 6505

89 31.7 6010

90 31.7 5505

91 31.7 5005

92 31.7 4520

93 31,7 4055

94 31.7 3550

95 31.7 3065

96 31.7 2525

97 31.7 2130

_83 31.7 9050

84 31.7 8570

85 31.7 8055

86 31.7 7535

87 31.7 7030

87 31,7 7030

88 31.7 6505

89 31.7 6010

90 31.7 5505

91 31.7 5005

92 31.7 4520

93 31.7 4055

94 31.7 3550

95 31.7 3065

96 31.7 2525

97 31.7 2130

X_lO0 35.7 4750

101 35.7 4750

102 35.7 4505

103 35.7 4010

104 35.7 3515

105 35.7 3005

106 35.7 2500

107 35.7 2170

'100 35.7 5790

101 35.7 4750

102 35.7 4505

I03 35.7 4010

104 35.7 3515

105 35.7 3005

106 35.7 2500

107 35.7 2170

_100 35.7 5790

101 35.7 4750

102 35,7 4505

103 35.7 4010

104 35.7 3515

105 35.7 3005

106 35.7 2500

107 35.7 2170

108 40 5670

109 40 5510

110 40 5010

111 40 4510

112 40 4000

113 40 3520

114 40 3015

115 40 2530

116 40 2160

TABLE B-1 (CONTINUED)

Power Gage No. of

Torque Coeff _N°" Peaks

25 1. 0505 1;01-4 0

15 .8858 B01-4 0

290 .9486 I;01-1 "_-

260 ,9484 B01-I 3

240 .991 881-1 3

210 .9909 1:131-1 1

280 .9758 8t31-1 3

155 .9814 881-1

135 1.0013 881-i 2

115 1.0167 881-1 2

95 1,016 BOI-1 0

75 .9835 801-1 2

65 1.0591 1101-I 2

45 •9566 JO 1 - I 0

35 .9982 BGI-I I

25 1.0505 801-I 0

15 •8858 BG 1 - I .9--.

290 .9486 801-6 5

260 . 9484 801-6 3

240 .991 801-6 4

210 .9909 801-6 6

180 ,9758 BG1-6 5

180 .9758 901-6 4

155 .9814 081-6 6

135 I.0013 BG1-6 3

115 1.0167 881-6 5

95 1.016 881-6 4

75 ,9835 801-6 i

65 I , 0591 BO I-6 2

45 .9566 881-6 I

35 .9982 881-6 I

25 1.0505 B0i-6 1

15 .8858 801-6 0

105 1 , 2468 801-4 2

105 1 , 2468 8G 1-4 2

90 1 • 1081 881-4 I

70 1 • 1663 881-4 1

50 1.0842 E_01-4 0

40 1.1868 801-4 0

30 1.286 001-4 0

18 1.0241 801-4 0

145 .3596 801-I 4

105 1.2468 861-1 2

90 1.1881 801--i 4

70 I• 1663 Bgl- I 2

50 1.0842 8Gi-! i

40 I• 1868 1101 - I 1

30 1 • 286 801-1 2

18 1.0241 B81-I i

145 .3596 801-6 4

105 1.2468 80i-6 4

90 1 • 1081 801-6 5

70 I • 1663 801-6 2

50 1.0842 801-6 3

40 1 • 1868 I_ 1-6 2

30 1 • 286 881"6 3

18 1. 0241 881-6 0

170 1,4167 801-4 4

160 1,4119 801-4 3

130 1.3876 BOI-4 2

110 1. 4489 801-4 2

85 1.4233 981-4 0

65 1,4055 881-4 0

45 1.3263 B01-4 1

35 1,4649 801-4 0

25 1 *4356 881-4 0

Spectral Frequencles(HZ)/VlbratorT Stress (psi)

152 570 252 672

238 585 244 582

134 922 238 693

126 550

11J 572 _ 644

108 937 216 1877

tO0 549 200 853

184 7_5 4_0 549

190 _34 396 705

390 947 442 501

386 596

454 718

286 536

268 502

254 6t2

440 863 454 2004 630 766

428 2050 626 859 890 870

134 910 434 931 624 651

126 672 426 871 432 806

116 948 426 910 430 922

118 806 424 903 610 606

108 819 216 879 418 694

202 548 414 868 846 818

184 601 408 1387 412 999

400 718 406 678 416 838

396 1485

398 1821 444 890

390 573

388 1073

382 529

410 974 684 649

676 777 680 855

680 830

672 599

888 708 898 920

878 831

438 883 624 1016

632 569 874 623

862 607

620 521 848 522

838 610 842 649

832 607

872 873

854 603

192 1140 204 1482 400 676 412 5418

192 1105 400 3252

190 1082 386 572 390 550 398 553

184 891 392 689

178 845

174 1338

170 1162 178 773

170 509

6 685 _824 1615 836 643

192 680 388 604 400 8247 826 635

188 757 386 1414 390 1453 426 525

392 1041 400 517

388 602 396 670 816 504

176 712 386 707

172 565 176 561 380 512

404 597 684 617 690 742 698 712

686 776 692 618 700 596

680 536 686 697

676 615 680 709

026 512

174 530

99 ÷
Run

No.

108

109

110

111

112

113

114

115

116

108

109

110

111

112

113

114

115

116

117

118

119

_120

t21

122

123

124

125

126

117

118

119

-120

121

122

123

124

125

126

117

t18

--120

121

122

123

124

125

126

127

128

129

130

131

132

133

134

135

127

128

129

130

Ref.

Blade

Angle

40

10

40

40

40

40

40

40

40

40

40

40

40

40

40

4O

40

40

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50,3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

50.3

60

&O

60

60

60

60

60

60

60

60

60

60

60

5670

5510

5010

4510

4000

3520

3015

2530

2160

5670

5510

5010

4510

4000

3520

3015

2530

2160

6225

6030

5520

5010

4505

4015

3520

3025

2520

2136

6225

6030

5520

5010

4505

4015

3520

3025

2520

2136

6225

6030

5520

5010

4505

4015

3520

3025

2520

2136

5280

5215

5000

4510

4025

3500

3045

2510

2150

5280

5215

5000

4510

TABLE B-1 (CONTINUED)

Power Gage No. of

Torque Coeff No. Peak.___ L

17o 1.416_ 801-1 5

160 1.4119 8G1-1 4

130 1. 3876 80 I- I 2

110 1. 4489 BG1 - 1 3

85 1. 4233 801-1 2

6_ I. 4055 BGI- I 2

45 I, 3263 881-1 I

35 I • 4649 8 tl I- 1 I

25 I. 4356 801 - I I

170 1.4167 801-6 6

160 1.4119 881-6 7

130 1 • 3876 801-6 6

110 1,4489 BD1-6 5

85 1. 4233 B81-6 2

65 I • 4055 881-6 3

45 1,3263 BG1-6 3

35 1.4649 BGl-6 3

25 1,4356 881-6 I

280 I •9358 B81-4 2

260 1.9157 B81-4 2

230 2. 0223 801 +4 2

190 2.028 861-4 2

155 2,0461 BG1-4 0

125 2,0775 861-4 2

95 2.0541 BGI-4 0

70 2,0495 BGI-4 0

50 2.1094 B81-4 0

35 2.0552 BGI-4 0

280 1.9358 BGI-I 6

260 1.9157 861-I 6

230 2.0223 801-1 6

190 2.028 801-1 8

155 2,0461 8Gi- 1 6

125 2.0775 801-1 3

9'5 2,0541 BG1-1 4

70 2,0495 801-1 1

50 2.1094 881-1 _.

35 2.0552 881-I I

280 1 •9358 BG I-6 7

260 1.9157 B01-6 7

230 2 •0223 BG 1-6 11

190 2,028 88l -_ 6

155 2,0461 Bnl-6 3

125 2 , 0775 BG I-6 2

95 2.0541 881-6 4

70 2 , 0495 BG1-6 2

50 2 , 1094 801-6 !

35 2,0552 801-6 1

290 2,7869 8G1-4 1

280 2,7583 801-4 1

260 2. 7863 881-4 1

220 2,8978 881-4 1

180 2,9767 881-4 0

135 2.9525 BGI-4 0

100 2 ,8895 BG 1-4 0

65 2 • 7641 881-4 0

50 2 • 8979 BG 1-4 0

290 2 . 7869 801-4 3

280 2 , 7583 801-4 2

260 2 • 7863 801-4 1

220 2,8978 881-4 I

Spectral Frequencles(HZ)/V[bratorv Stress (gsl)

94 620 188 5:7 202 1934 402 2309 406 2215

188 726 198 1034 398 1812 406 1132

19o 1184 4o2 1o68

19o 1282 :o2 675 396 811

180 2003 184 2435

160 515 180 1090

176 2839

170 981

I_ 993

94 599 202 656 402 7650 420 534 836 649 840 580

92 566 200 811 396 4004 406 2919 414 1571 834 791

842 902

84 606 194 663 386 558 402 2968 418 560 830 764

188 978 382 511 392 1115 398 1136 826 746

182 1945 392 727

176 705 182 547 390 615

176 1955 378 519 384 782

170 693 378 510 386 625

168 663

702 1012 710 810

690 749 704 1051

690 542 696 800

682 576 688 738

182 929 676 557

104 1383 184 660 202 4431 208 2753 404 858 412 730

100 1273 182 676 188 918 _._02 5513 212 767 406 955

92 1432 184 1569 196 2464 21 _2 533 402 607 406 614

84 908 158 512 162 590 168 926 192 2782 200 596

230 743 398 598

74 773 186 2145 194 276 200 771 206 839 394 537

180 3949 184 3686 392 617

[62 I066 174 I448 178 1840 388 541

174 1836,

168 738 172 787

168 820

202 2784 384 511 404 2026 430 516 618 7_7 840 _50

848 747

100 536 202 2333 388 641 406 2135 616 825 838 744

846 897

174 506 184 827 194 934 200 737 386 524 400 1368

404 1333 410 706 616 627 834 925 838 965

192 994 390 888 394 1032 404 722 8_8 686 83_ 658

186 672 394 1122 826 563

182 3305 390 1232

162 635 176 790 388 754 816 546

172 824 384 839

|70 913

170 549

684 502

686 520

684 565

678 570

176 901 190 1253 194 972

174 512 190 1317

188 1340

184 2268

ORIGINAL P/iGE |:_

OF POOR QUALITY

I00
oF POOR

Run

No.

131

132

133

134

135

127

128

129

130

131

132

133

134

135

135

137

138

139

140

141

162

136

137

138

139

140

141

142

136

137

138

139

140

141

142

143

144

145

146

147

148

149

143

144

145

146

147

148

149

143

144

145

146

147

140

149

150

151

152

153

154

155

156

157

150

Ref.

Blade

Angle

60

60

60

60

60

60

60

60

60

60

6O

60

60

6O

60

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

69.9

B0

B0

80

80

80

B0

B0

8O

80

80

80

80

80

80

80

80

80

80

80

80

80

34

34

34

34

34

34

34

34

34

Re___M

4025

3500

3045

2510

2150

5280

5215

5000

4510

4025

3500

3045

2510

2150

2150

4500

4000

3510

3025

2510

2130

4785

4500

4000

3510

3025

2510

2130

4785

4500

4000

3510

3025

2510

2130

5000

4520

4020

3530

3025

2510

2150

5000

4520

4020

3530

3025

2510

2150

5000

4520

4020

3530

3025

2510

2150

5700

5010

4525

4015

3515

3015

2520

2167

5700

180

135

100

65

50

290

280

260

220

189

135

100

65

50

50

260

205

160

118

83

58

290

260

205

160

• 118

83

58

290

260

205

160

118

83

58

28O

230

185

140

100

70

48

280

230

185

140

100

70

48

280

230

185

140

100

70

48

115

IO0

78

62

45

32

22

15

115

TABLE B-1 (CONTINUED)

Powe r

Coeff

2.9767

2.9525

2.8895

2.7641

2.8979

2.7869

2.7503

2.7863

2.8970

2.9767

2.9525

2.8895

2.7641

2,8979

2.8979

3.4399

3.4326

3.4793

3.4548

3.5296

3.425

3.3933

3.4399

3.4326

3.4793

3.4548

3.5296

3.425

3.3933

3.4399

3.4326

3.4793

3.4548

3.5296

3,425

3.0006

3.0161

3.067

3.01

2.9278

2.9767

2.782

3.0006

3.0161

3.067

3.01

2.9278

2.9767

2.782

3.0006

3.0161

3.067

3.01

2.9278

2.9767

2.782

.9483

1.0674

1.0206

1.0304

.9758

.9431

.9281

.8558

.9483

Gage No. oF

No. Peak._Ls Spectral Frequencles(HZ)/Vlbratory Stress (l_si)

_G l.-e L 1_10 1535

_81-4 1 176 2022

881-4 1 174 1289

881-4 1 168 818

901-4 i 168 1139

201-6 6 194 792 396 843 402 579 616 618 622 528 832 734

N|-6 4 190 823 394 725 614 564 830 601

801-6 4 188 1155 396 941 612 694 826 678

101-6 3 184 1614 390 1019 452 510

!01-6 2 llO 1378 386 974

801-6 2 176 1376 384 962

801-6 2 172 726 382 735

801-6 I 170 518

BGI-6 1 168 626

801-4 1 676 510

801-4 0

8GI-4 0

BG1-4 0

881-4 0

801-4 0

8G1-4 0

881-I I 184 1640

BGI-1 1 182 1630

BGI-t 1 180 1022

BGI-I I 176 1331

881-1 1 170 869

Br:l-I I 170 686

881-I 1 16e 571

80i-6 4 184 887 394 1203 444 571 824 550

801-6 4 180 891 390 981 394 917 820 570

801-6 2 176 591 388 1125

BOI-6 2 176 885 384 851

80i-6 2 i72 590 382 912

801-6 I 376 532

801-6 0

801-4 I 672 760

801-4 1 664 544

BG1-4 0

801-4 0

801-4 0

BG1-4 0

801-4 0

861-I 4 166 1576 184 3451 200 670 392 947

801-1 3 150 701 182 3280 392 620

BGI-1 3 134 587 176 3306 386 820

801-1 3 118 548 172 2355 384 524

BGI-I 1 170 1472

881-1 2 166 1104 172 1057

8G1-1 1 168 911

801-6 9 166 1066 182 2046 372 539 392 2299 414 511 430 541

436 710 610 702 822 945

BGI-6 5 168 500 180 2341 390 1652 398 766 820 706

B81-6 2 176 2054 384 1433

BDI-6 2 172 1274 384 1120

801-6 2 168 1033 378 730

801-6 1 166 810

801-6 I 166 971

801-4 2 410 1013 684 954

801-4 2 682 805 686 796

BQI-4 1 678 640

BQI-4 1 672 524

801-4 o

BQi -4 0

801-4 b

801-4 0

901-1 3 204 1064 400 561 410 4352

IOI
Run

No.

i

151

152

153

154

155

156

157

15o

151

152

153

154

155

156

157

158

159

16o

161

162

163

i64

t65

L66

162

168

tb9

i70

171

[72

158

159

160

|61

--162

163

164

165

166

167

168

169

170

171

172

158

159

160

--161

162

163

164

165

166

167

168

169

170

171

172

176

177

Ref.

Blade

Angle

34

34

34

34

34

34

34

34

34

34

34

34

34

34

32.7

32.7

32.7

32.7

32.7

32.7

32.7

32.7

32,7

32.7

32.7

32.7

32.7

32.7

32.7

32.7

32.7

32.7

32.7

32.7

32.7

32,7

32,7

32.7

32,7

32.7

32.7

32.7

32.7

32.2

32,7

32.7

32.7

32.7

32.7

32.7

32.7

32,7

32.7

32.7

32.7

32.7

32.7

32.7

33.7

44.9

44,9

TABLE B-1 (CONTINUED)

Power Gage No. of

RPM Torque Coeff No. Peaks 100 I.0""_74 _ot-1 2

4525 7s 1.o2o+ 901-1 3

4015 62 1.0304 801-i 4

3515 45 .9758 981-1 1

3015 32 .9431 BOI-I 2

2520 22 .9281 901-1 1

2167 15 .8558 881-I 1

5700 115 .9483 BGI-6 4

5010 100 1.0674 801-6 3

4525 78 1,0206 981-6 3

4015 62 1.0304 881-6 I

3515 45 .9758 B01-6 1

3015 32 .9431 801-6 1

2520 22 .92gl 981-6 1

2167 15 .8558 901-6 0

9025 280 .92i 981-4 4

0517 265 .9787 801-4 2

8030 230 .9556 B01-4 2

7530 210 ,9922 961-4 2

7020 175 .9514 901-4 2

6585 155 ,9577 801-4 1

6025 135 .9963 901-4 2

5530 110 ,9637 BG1-4 2

5040 100 1.0547 981-4 2

4510 75 .9879 B01-4 I

3990 60 1.0097 901-4 1

3470 45 1.0013 981-4 0

3055 35 1.0047 981-4 0

2485 20 .8677 BGI-4 0

2110 15 .9026 901-4 0

9025 290 ,921 801-1 4

8517 265 ,9787 881-1 5

8030 230 .9556 901-1 4

7530 210 .9922 9GI-1 3

7020 175 .9514 881-1 5

&595 155 .9577 961-I 3

6025 135 .9963 BGl-I 4

5530 110 ,9637 9GI-I 3

5040 100 1.0547 901-1 2

4510 75 ,9979 861-I 2

3990 60 1,0097 BG1-1 3

3470 45 1.0013 BGI-1 o

3055 35 1.0047 881-1 2

2485 20 .8677 901-I I

2110 15 .9026 BGI-I I

9025 280 .921 981-6 9

8517 265 .9787 BG1-6 7

8030 230 .9556 B01-6 9

7530 210 .9922 BG1-6 7

7020 175 .9514 901-6 5

6585 155 .9577 801-6 6

6025 135 .9963 B01-6 6

5530 110 .9637 BGI-6 4

5040 I00 1,0547 _81-6 2

4510 75 .9879 881+6 3

3990 60 1.0097 BGI-6 2

3470 45 1.0013 BGI-6 I

3055 35 1.0047 BGI-6 2

5495 20 .8677 B01-6 0

3110 15 .9026 881-6 0

6800 280 1.6223 881-4 4

6520 270 1.7016 981-4 3

Spectral Frequencies(HZ)/VlbratoryStress (pst)

196 564 408 578

190 724 388 789 396 524

178 583 188 744 392 1084 398 543

180 664

174 655 178 614

170 949

168 699

410 12449 822 694 "834 672 840 570

404 1532 420 553 834 708

390 2932 400 1026 824 705

392 2655

388 875

386 700

172 578

628 516 632 507 746 867 754 ?62

740 1252 746 817

728 770 732 805

720 636 726 629

696 514 714 959

706 765

690 661 696 772

684 547 690 753

680 573 688 537

674 648

672 673

150 802 246 541 250 573 442 510

142 581 236 644 240 754 248 813 284 500

134 655 222 568 232 616 238 790

230 827 252 736 426 538

220 540 226 569 234 712 416 529 420 514

110 738 214 1003 220 1051

100 703 202 1201 208 850 410 505

184 996 204 747 404 1147

194 775 404 509

192 665 196 637

182 688 396 1338 440 518

176 870 384 521

174 623

170 691

150 618 300 527 430 553 438 764 448 _44 62R 1437

892 1064 896 948 908 615

142 656 428 1160 438 859 626 833 742 754 882 94,5

888 748

134 638 430 792 440 617 446 512 452 518 622 772

870 686 878 1105 888 718

416 507 426 1126 440 920 622 1115 856 550 864 762

872 716

414 1396 420 1579 434 550 620 942 858 942

110 753 220 764 414 911 612 630 620 750 852 565

200 612 408 1080 416 1313 422 793 614 607 842 818

184 647 406 2063 610 519 836 618

404 1022 830 610

384 839 394 1060 402 641

396 3191 442 1490

392 678

386 836 392 552

704 720 710 690 716 1537 732 503

706 690 710 926 720 626

ORiGLNAL pF._Z IS

OF POOR QUALITY

102
O_iGINAL Pt_GE L-3

OF POOR Q;_ALITY

Run

No.

17ff

179

180

181

182

183

184

185

1 B6

176

177

178

179

180

181

182

183

184

185

186

176

177

Ref.

Blade

Angle

44.9

44.9

44°9

44.9

44.9

44.9

44.9

44,9

44.9

44,9

44,9

44.9

44,9

44.9

44.9

44.9

44.9

44.9

44.?

44.9

44*9

44.9

TABLE B-1 (CONTINUED)

Power Gage No. of

RPH Torque Coeff No. Peaks

604"S 2_ 1,7_J9 SOt-4 2

5500 195 1.727 861-4 3

5000 165 1 • 7682 BG 1-4 1

4500 135 1.7861 801-4 1

4000 105 1. 7582- 801-4 1

3495 80 1. 7546 801-4 o

3018 60 1,7648 801-4 0

2505 40 1 , 7078 BOl-4 0

2157 28 1.6123 DG1-4 0

6800 280 I • 6223 BG1 - i 12

6520 270 1 • 7016 801-1 B

6045 235 1 . 7229 B81 - 1 6

5500 195 J • 727 BG1-1 9

5000 165 1 . 7682 881-1 7

4500 L35 I • 7861 BGI-1 5

4000 L05 1• 7582 BG 1-1 Z

3495 80 1 , 7546 881 - 1 3

3018 60 1 , 7648 881 - 1 1

2505 40 I . 707B BG1-1 I

2157 28 1.6123 881-1 1

6800 2B0 L• 6223 BG 1-6 13

6520 270 1,7016 801-6 14

178 44.9 6045 235 1.7229 8G1-6 10

179 44.9 5500 195 1,727 BOI-6 9

180 44.9 5000 165 1.7682 B01-6 7

181 44,9 4500 135 1.7861 B01-6 5

182 44.9 4000 105 1.7582 BGL-6 3

I83 44.9 3495 80 1*7546 BG1-6 2

184 44.9 3018 60 1.7648 881-6 3

185 44.9 2505 40 1.7078 B01-6 2

186 44.9 2157 28 1.6123 B81-6 1

189 38 5725 165 1.3487 B01-4 4

1_0 38 5_20 150 1.3189 B01-4 2

191 38 5020 125 1.3289 BG1-4 2

192 38 4518 100 1.3125 881-4 2

193 38 3995 BO 1o3429 881-4 1

194 3B 3520 60 1.2974 BGI-4 _ t

195 38 3000 45 1.3396 BQI-4 0

L96 38 2510 33 1.3922 BGI-4 0

197 3B 2180 21 1,1839 BGI-4 b

189 38 5725 165 1.3487 B01-1 S

190 38 5520 L50 1.3189 881-I 4

191 38 5020 L25 1.3289 BGI-L 3

192 38 4518 LO0 1.3125 801-1 2

193 38 3995 BO 1.3429 B01-1 I

194 38 3520 60 1,2974 BG1-1 1

195 38 3000 45 1,3396 801-1 1

196 38 2510 33 1,392_ 881-1 I

197 38 2180 21 1.1839 881-1 I

IB9 38 5725 165 1.3487 801-6 4

190 38 5520 150 L.3189 BGI-6 3

191 38 5020 125 1,3289 801-6 7

192 38 4518 100 1,3125 801-6 6

193 38 3995 80 1,3429 801-6 4

S,pectral Freq,uenctes(HZ)/Vibratory Stress (psi)

700 1032 710 645

68B 664 696 754 706 535

688 991

682 709

184 670

114 1349 194 677 198 749 210 5158 226 912 318 582

396 610 404 1001 412 994 420 632 428 511 718 533

110 971 190 68 "_ 204 3497 208 3883 218 1804 390 629

412 1369 422 513

100 1668 190 1609 202 2028 208 1458 220 511 410 1173

P2 915 170 502 176 570 I84 897 194 1945 204 1031

254 542 396 584 406 761

166 607 182 546 190 1593 200 669 234 800 394 903

402 "727

74 685 188 1062 200 759 206 1055 394 599

66 6130 182 7468

160 1106 170 534 180 914

174 2515

170 1549

168 1402

212 1933 220 619 384 509 390 725 400 2340 "406 1662

414 2519 420 2121 432 512 624 1640 696 525 716 619

856 1042

204 3036 216 839 368 551 382 896 402 997 408 1640

416 1568 424 715 434 640 616 1109 624 1001 690 649

844 661 850 1100

100 5_J8 L88 770 196 517 202 1058 392 598 404 1809

412 2464 620 745 838 628 848 897

196 1046 202 565 388 727 398 1474 406 1407 414 1125

422 525 616 660 838 907

194 581 388 549 396 L906 406 664 456 543 830 541

834 617

184 817 206 517 386 523 396 1444 828 652

184 3059 390 665 394 712

178 866 390 689

172 1330 380 832 386 1418

170 757 380 725

168 856

408 672 690 913 694 717 698 573

690 708 698 599

682 867 690 667

676 655 682 644

676 743

670 610

190 606 196 762 204 785 210 582 408 3816

184 613 198 1185 394 757 402 2583

84 551 194 1336 402 748

188 1346 398 702

18o 1728

178 933

174 2263

172 1049

168 746

190 532 400 1996 408 9162 840 774

200 640 400 7823 838 840

84 546 194 650 392 _98 40_ 2091 414 999 610 507

832 1068

190 744 384 850 390 971 396 1096 446 515 826 593

180 1153 184 906 390 951 3_:_ _12

103
Ref.

Blade

Run Angle

No.....= _ RP._M_M

194 38 3520

195 38 3000

196 38 2510

197 38 2180

198 -10 9010

199 -10 8560

200 -10 8020

201 -10 7525

203 -10 6500

204 -10 5985

205 -10 5540

2O6 -10 5010

207 -10 4530

208 -10 3990

209 -10 3510

210 -10 3005

211 -10 2100

198 -10 9010

199 -I0 8560

200 -10 g020

20[ -I0 7525

203 410 6500

204 -10 5985

205 -1o 5540

206 -10 5010

207 10 4530

208 -10 3990

209 -lO 3510

210 -10 3005

211 tO 2100

L98 -t0 9010

199 -I0 8560

200 -10 B020

201 -10 7525

203 -10 6500

204 -10 5985

205 -1o 5540

206 . -10 5010

207 -i0 4530

208 -10 3990

209 -tO 35t0

210 -tO 3005

211 -to 2100

Torque

60

45

33

21

12

11

10

9

8

7

6

5

4

3

3

2

0

12

11

lO

9

8

7

6

5

4

3

3

2

o

12

11

lO

9

8

7

4

3

3

2

0

Powe r

Coeff

1.2974

1 • 3396

1,3922

I• 1839

• 0396

.0402

.0417

.0426

.0507

• 0524

.0524

.0534

.0522

.0505

.0652

• 0593

0

.0396

.0402

.0417

.0426

•0507

.0524

.0524

.0534

.0522

.0505

• 0652

.0593

0

.0396

• 0402

.0417

•0426

.0507

• 0524

.0524

• 0534

,0522

• 0505

.065_

.0593

0

TABLE B-1

Gage

No.

801-6

BOI-6

gOI-6

801-6

801-4

BGI-4

B01-4

801-4

Bnl-4

801-4

B01-4

801-4

801-4

BIBI-4

801-4

801-4

801-4

BGI-i

B81-t

801-1

BGI -t

BGllt

BG1 -t

BGI-I

BGI-I

BGI-I

801 -I

BG1-1

BG| -I

BO1-1

801-6

801-6

BGI-6

BO1-6

BG1-6

BG1-6

B01-6

801 - 6

BGI-6

BG1-6

BGi -6

BGI 6

B01-6

(CONTINUED)

ORIGINAL PAGE IS

.OF..POOR QUALITY

No. of

Peaks

2

2

2

1

0

0

0

0

0

0

0

0

0

0

0

0

0

1

1

1

1

0

0

0

0

0

0

o

0

o

1

]

o

0

0

0

0

o

0

0

0

0

0

Spectral Freq_encles(HZ)/Vlbratory Stress

17B 632 390 840

174 1312 386 738

172 780 382 527

170 606

150 668

142 590

134 520

126 507

150 528

144 517

(Dsl)

104
APPENDIX C

SR-3 CAMPBELL DIAGRAMS

Thls appendix contains Campbell diagrams from the zero forward speed SR-3

model Prop-Fan tests conducted at UTRC, given In terms of response frequency

vs. RPM for various blade angles. These Campbell diagrams were generated

from the data obtained from the spectral analyses data using computerized

peak-plcklng routines. These are the same data that are tabulated in

Appendix B and were plotted automatically by computer. Shown are plots of

frequency versus rotational speed. A plot Is generated for each blade

angle. Moda] response frequencies are evldent at the higher blade angles.

FIGURE NO.

C-1

C-2

C-3

C-4

C-5

C-6

C-7

C-8

REFERENCE BLADES-ANGLES, DEG

-I0.0, 12.0

15.9, 19.9

23.6 27.6

31.7, 32.7

34.0, 35.7

38.0. 40.0

60.0, 69.9

80.0

105/'106

r- ORIGINAL PAGE IS

OF POOR OUALITY

[ 03

1.00

SR-3 _'MTIC l"lb'l"

UTRC UZHD TUPCqCf.,

/Y/1/ , /

.- ///////

e. 8e /=_

..,._ .. /U/_/ / /

O, ,- _...,.._4P i" /, //3// _/

0. SO

• _--_P

.. //_ _--__ _.--_

_._ _ , , ,

0.04t 41.1_ 0._0 e.3_ O.,le e.5_ 0.60 O.?e O 80 0._ 1.00

FIGURE c-i. SR-3 PROP-FAN STATIC TESTS, CAMPBELL DIAGRAMS FOR

REFERENCE BLADE ANGLES OF-10 DEG'S AND 12.0 DEG'S

lO7
D_A_ P._GE IS

OF POOR q_T._ -_v

Slq-3 IT_TZ¢ TEST

[ 03 U'm¢ WZND

1,iN) , , , ,

I

'z_p111_elb_ /p /-ab /P ,._ /_// / / /

... /Z_//,/ _j

/,U,'/_//f

/////'//./" ....

- /

/,'///////

#.N

--,---' , , , 1, ,

#,_ e,le #.|e Q,3e B.4o o,_ o,_ e,vo o.8e o._ 1.oo

e.?e F

R

£ I).6e

o

u

E e.se H

¢

¥ e,,le

H

z t.ll

sR-3 ST_TZC TEST

E 03 U'rll¢ WZHD

I

Z/l//

,.. Z U.'// / .<._°'-

• l//i'll / _,,

[ 0.80 _. i .. //Y///',// _/

i .. ///,V/'/./.. -/ .--"

i ¢,'//_//J,/-_ J

e.le--_--.

""" _/_<_---__-- ,1

,.,,"." _ , _ , , , _

FIGURE C-2. SR-3 PROP-FAN STATIC TESTS, CAMPBELL DIAGRAMS FOR

REFERENCE BLADE ANGLES OF 15.9 DEG'S AND 19.9 DEG'S

108
f

ORIGINAL PAGE tS

OF POOR QUALITY

F

R

E

0

U

i

N

O

V

H

Z

Sl-3 trA'riO TElrl'

E e3 u'mo Wield

1. ee --T ..... , , ' ' 12 /liP. / _ / ,/

..- ////////

e lit I"P

ll.lw i

0. le __i_--_,---- . , , ,,

l.it t,_ t.tl I.?e I.il t._ 1,_

Eli,4

0.30 o.4e O.SO

sP_D-mm

FIGURE C-3. SR-3 PROP-FAN STATIC TESTS, CAMPBELL DIAGRAMS FOR

REFERENCE BLADE ANGLES OF 23.6 DEG'S AND 27.6 DEG'S

109
ORIGINAILPACE

OE Poor Qu_

$11-3 _'RgZC 't'Irr

i _I UTRC WZND

/ I / / u,,.

!

41, F I, //,/,,/./:/;,.y. _

_"'. /,'9"///P.'/" _/ -

O.SO ...

Y lt.,4t

tll.llll IP.ilP Ill._lt tt.31t li._ 0.110 il.tiO ll.?tl Ill.ill_ i.illt 1.Oil

SPEED-_P_I E

[ i3

l. litt

SR-3 _rATz¢ 'rEsw

WIND "rkx,i,_

,, , , , , la_tl#litl tl D /P "_ ,

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REFERENCE BLADE ANGLES OF 80.0 DEG'S

114