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NASA CR.179528

DYNAMIC RESPONSE AND STABILITY OF A

COMPOSITE PROP-FAN MODEL (NASA-CB- 179528) DYIAMIC EESECNSE AND ~ a 7 - 2 1956 ZTABILXTY OF A C C P P O S I T E ERCE-PA6 40DEL F i n a l Report LHariltcn Standard) 92 p A v a i l : YTXS HC L I C S / t l F A01 C S C L 21E Unclas

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,j G3/07 0072745

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BY Arthur F. Smith

and Bennett M. Brooks

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HAMILTON STANDARD DIVISION UNITED TECHNOLOGIES CORl?ORATt,ON I

WINDSOR LOCKS, CONhlECTlCUT 06696

i

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~

October, 1986

National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio 441 35

% f Contract NAS9-24088 9i

https://ntrs.nasa.gov/search.jsp?R=19870012523 2020-03-23T21:26:15+00:00Z

NASA C R-179528

1 DYNAMIC RESPONSE AND STABILITY OF A

COMPOSITE PROP-FAN MODEL

Arthur F. Smith and

Bennett M. Brooks

HAMILTON STANDARD DIVISION UNITED TECHNOLOGIES CORPORATION WINDSOR LOCKS, CONNECTICUT 06096

October, 1986

National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio 441 35

Contract NAS3-24088

12. Government Accession No. 1. Reporl No. 3. Reclpient'a Catalog NO.

Dynamic Response and S t a b i l i t y o f a Composite

Prop-Fan Model

CR- 179528 4. Title and Subtitle

October, 1986 6. Performing Organization Code

5. Reporl Date

7. Author@)

* + A. F. Smith and 8. M. Brooks

8. Performlng Organization Report No.

0. Perfoming Organization Name and Address

Hami 1 t o n Standard D i v i s i o n Un i ted Technologies Corporat ion Windsor Locks, CT 06096

2. Sponsoring Agency Name and Address

Nat iona l Aeronaut ics and Space Admin i s t ra t i on Washington, D. C. 20546

11. Contract or Grant No.

NAS3- 24088 13. Type of Reporl and Period Covered

Cont rac tor

I

5. Supplementary Notes

Advanced Turboprop Prctpel ler Aeroel a s t i c Test Composite S t r u c t u r a l Response Energy E f f i c i e n t Wind Tunnel Test

Prop- Fan

F i n a l Report. P r o j e c t Technical Monitor, M r . 0. Mehmed, NASA-Lewis Research Center, Cleveland OH 44135.

* ,Nrow w i t h : H i rock Corporation, 1 Main Road, G r a n v i l l e , MA 01034 + Now w i t h : Ko1l;;lorgen Corporat ion, 347 King S t ree t , i iorti iampton, MA 01060

6. Abstract

Results are presented for blade response and stability during wind tunnel tests of a 62.2 cm (24.5 in) diameter model of a Prop-Fan, advanced turboprop, with swept ,graphite/epoxy composite blades. Measurements of dynamic response were made with the rotor mounted on an isolated nacelle, with varying tilt for non-uniform inflow, at flow speeds from 0.36 to 0.9 Mach number.

The blade displayed no instabilities over the operating range tested, up to 0.9 Mach number and 10000 RPM.

Measurements are compared with those for other Prop-Fan models of both solid metal and graphite composite construction. composite blade had less response than an unswept (straight) composite blade. blades.

The swept

Composite blades had more response than metal

Unl i m i t e d .

Measurements are compared with theoretically based predictions. The 1-P blade response was significantly overpredicted using unimproved methods and somewhat overpredicted using improved methods.

9. Security Classif. (of thle report) . M. security C\.ssif. (of thir Paw) 21. No. of piger

U n c l a s s i f i e d 94 Uncl ass i f i ed

Unexpectedly h i g h 2-P strain levels were measured and suggest the presence o f non-linear effects on blade response.

22. Prlce'

7. Key Words (Suggested by Author@)) 118. Dlrtrlbution Statement

'For sale by the National Technical Informallon Servlce. Sprlngfield. Virginia 22161

FOREWORD

All of the testing reported herein was performed by NASA-Lewis personnel in the 8 x 6 wind tunnel at NASA-Lewis. The data were reduced, analyzed and reported by personnel from Hamilton Standard, a division of United Technologies Corporation.

I I

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Mr. Oral Mehmed was the NASA Technical Monitor for this project. At Hamilton Standard, Mr. Donald Marshall performed the data reduction, Mr. Prem Bansal I

I performed the theoretical predictions with assistance from Mr. Peter Arseneaux, and Mr. Arthur F. Smith conducted the data analysis and correlation with predictions. Mr. Bennett M. Brooks was the Hamilton Standard Project Manager.

I I

This work was accomplished under contract NASA-24088 for the NASA Lewis Research Center in Cleveland, Ohio. ,

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TABLE OF CONTENTS

Page

1 . 0

ABSTRACT ................................................. FOREWORD ................................................. TABLE OF CONTENTS ........................................ SUMMARY .................................................. SYMBOLS AND ABBREVIATIONS ................................ INTRODUCTION .............................................

2.0 DESCRIPTION OF THE EXPERIMENTAL PROGRAM

2.1 Test Model ............................................ 2.2 Wind Tunnel Facility .................................. 2.3 Model Instrumentation ................................. 2.4 Test Procedures ....................................... 2.5 Test Conditions ....................................... 2.6 Data Reduction ........................................

3.0 ANALYTICAL TECHNIQUES

3.1 Approach .............................................. 3.2 Calculated Modes and Frequencies ...................... 3.3 Calculated Vibratory Strains ..........................

4.0 TEST DATA EVALUATION AND COMPARISON WITH CALCULATIONS

4.1 4.2 4.3 4.4 4.5 4.6 4.7 4.8 4.9 4.10

Flutter Clearance ..................................... Total Strain Measurements ............................. Spectral Analysis ..................................... Modal Frequencies and Comparison to Predictions ...... P-Order Vibratory Strains ............................. Shaft Tilt and the Excitation Factor .................. Power Coefficient ..................................... Strain Sensitivity vs . Power Coefficient .............. Comparison of l-P Measurements to Predictions ........ Comparison to other Prop-Fan Models ...................

5.0 CONCLUSIONS .............................................. 6.0 RECOWNDATONS ........................................... 7.0 REFERENCES ...............................................

TABLES ................................................... FIGURES .................................................. APPENDIX I - Peak Total Strain Tabulation ................ APPENDIX I1 - P-order Strain Tabulation ..................

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V

1

3

7

7 7 8

11

11 11 12

14 15 16 16 17

19

21

23

25 35 67 75

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SUMMARY

TEST

The tests were conducted, in the NASA-Lewis Research Center 8 x 6 foot wind

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High speed dynamic response and stability tests were conducted on a model Prop-Fan, with swept composite blades, which was found to be structurally adequate over the entire operating range. I

DATA ANALYSIS and CORRELATION to CALCULATIONS

Blade vibratory strain gage test data were reduced and analyzed to determine response and stability trends for variations of operating parameters. Non-dimensionalized blade strain sensitivities are presented as a function of rotor power coefficient.

Calculations of blade response were made using lifting line aerodynamic and finite element structural methodologies. data. previously tested, Prop-Fan models of both solid titanium and graphite composite construction.

The calculations are compared to test Also, data for the SR-3C-3 model are compared to data for other,

CONCLUSIONS

The SR-3C-3 model was structurally adequate over the range of operation, demonstrating the success of composite structural tailoring in preventing instability.

The swept composite blade had less response than the straight composite blade . The trends of 1-P blade response were well defined using non-dimensional parameters.

The composite blades were more strain sensitive than the metal blades.

The SR-3C-3 1-P response was significantly overpredicted using unimproved methods. Improved finite element calculation methods reduced the amount of 1-P overprediction.

vii

SUMMARY (continued)

6 ) High measured 2-P strain levels suggest the presence of undetermined non-linear effects on blade response.

RECOMMENDATIONS

1) The improved finite element prediction method should be confirmed by additional 1-P calculations.

2) Existing test data for other Prop-Fan models should be reviewed to determine the extent of non-linear effects on blade response.

3) Non-linear effects should be included in future improvements t o the blade response calculation method.

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SYMBOLS AND ABBREVIATIONS

1 .o 1% x 3 dx

0.2

AF Blade A c t i v i t y Factor = 100,000

16

b Blade Sect ion Chord Width, m

c1 Blade Sect ion Design L i f t C o e f f i c i e n t

CP Power c o e f f i c i e n t = 2nQ/pn2D5 = I T ~ Q / , , ~ P V : I p D 3

D Rotor Diameter, m I

, EF E x c i t a t i o n Factor = I (Veq /348 l2

N Rotor Speed, RPM

n Rotor Speed, r e v o l u t ions/sec

0 Rotor Torque, N-m

SHP S h a f t Horsepower

V e q

V T True airspeed, knots

V T I P Blade T i p r o t a t i o n a l speed, m / s = nnD

X Non-Dimensional Blade Radius

B R E F . Reference Blade Angle deg

0 . 7 5

E M i c ro -S t ra i n

P A i r Densi ty , kg/m3

P o

I Prop-Fan s h a f t t i l t , degrees

1P Frequency = one per p r o p e l l e r r e v o l u t i o n , Hz

nP Frequency = n per p r o p e l l e r r e v o l u t i o n , Hz

Equ iva len t a i r speed = V T Jplp,,, knots

Blade Angle a t 3 /4 Radius = D R E F - 0.9, deg

A i r Densi ty, Standard Sea Level = 1.2250 kglm'

S I u n i t s of measurement used throughout unless s p e c i f i e d otherwise.

ix/x

1.0 INTRODUCTION

The efficiency advantage offered by the Prop-Fan has been demonstrated by extensive testing, see Reference 1. can achieve up to a 30 percent improvement in efficiency, over conventional means of propulsion. Since 1976, many experimental test programs have been conducted with model Prop-Fans. various model Prop-Fans were subjected to a complete range of operating conditions, including extremes in loading, such that the structural integrity of the blades as well as the advantages for rotor performance were demonstrated. The structural testing has included investigations of blade vibratory response due to angular inflow, unstalled flutter, buffeting, stall flutter, and vibratory response to a non-uniform flow field due to installation on an aircraft (wing/nacelle/f uselage) model.

The Prop-Fan models tested previously (see References 2, 3, and 4) of solid steel or titanium, and performed very well in view of the rigorous structural testing to which they were subjected. It is recognized that for full scale Prop-Fan applications, solid metal blades would be too heavy. Therefore, testing of Prop-Fan models constructed of lighter weight composite materials has begun. The objective of these tests is to demonstrate the efficiency advantage expected for Prop-Fans, while maintaining structural integrity in a design free of instabilities.

It is well established that the Prop-Fan

During the course of these test programs, the

were made

As part of the continuing studies of Prop-Fan structural stability and blade dynamic response, an 8-bladed model, designated the SR-3C-3, was designed by NASA-Lewis, with Hamilton Standard support, and fabricated by NASA-Ames. The SR-3C-3 Angular inflow and unstalled stability tests, with the Prop-Fan model mounted on an isolated nacelle, were conducted in the NASA-Lewis 8 x 6 foot wind tunnel, at free stream Mach numbers of 0.36 to 0.90. These tests were conducted during the period of July 11 through 19, 1983 providing test support under its own funding. Hamilton Standard analyzed the data acquired during these tests.

blade is a solid composite design of carbon fiber in an epoxy matrix.

with Hamilton Standard Then, under contract NAS3-24088,

This report summarizes the results of the SR-3C-3 Prop-Fan model dynamic response and stability investigation. strain data with operating conditions. strain, P-order strain and frequency spectra were analyzed. dynamic responses were predicted, using theoretically based calculation procedures for comparison to test results. The predicted structural mode shapes and frequencies were provided by NASA-Lewis, while the airloads and structural responses were calculated by Hamilton Standard, using a NASA-Lewis supplied MSC/NASTRAN finite element model of the composite Prop-Fan. comparisons were used to evaluate the ability of the theoretical method to predict blade loading and response, in order to verify the method's usefulness as a Prop-Fan design tool.

Included are trends of measured blade Total vibratory strain, modal vibratory

In addition, 1-P

The

2.0 DESCRIPTION OF THE EXPERIMENTAL PROGRAM

The tests described in this report were conducted using the SR-3C-3 8-way Prop-Fan model, mounted on an isolated nacelle, in the NASA-Lewis Research Center 2.44 x 1.83 m (8 x 6 ft) wind tunnel. The primary purpose of these tests was to determine the effects of yawed flow, at Mach numbers from 0.36 to 0.84, on the blade's vibratory response. stability of this configuration was investigated up to 0.9 Mach number.

In addition, the aeroelastic

2.1 Test Model

The SR-3C-3 Prop-Fan model is nominally 62.2 cm (24.5 in.) in diameter. The Prop-Fan concept incorporates thin airfoils (2 percent thick at the tip) and swept blades to achieve high aerodynamic efficiency with low noise generation. The SR-3C-3 an earlier Hamilton Standard design. characteristics of this blade can be found in Reference 3. Table I is a summary of the overall design parameters for the SR-3C-3 model.

geometric shape is identical to that of the SR-3 model, which is A description of the geometric

The SR-3C-3 model blades were built by NASA-Ames and the hub was built by Hamilton Standard. epoxy matrix. construction. Reference 5. degrees. The fiber ply orientation was chosen to provide the blade with similar structural vibratory response frequencies to those for the metal (titanium) SR-3 model, and to allow the model to be free of instabilities.

The blades are made of carbon fiber cloth layers in an This model is one of a series of SR-3C blades of similar A more complete description of the blade series is given in In the SR-3C-3 model, the carbon fibers are oriented at 245

2.2 Wind Tunnel Facility

Figure 1 shows the SR-3C-3 model Prop-Fan rotor installed in the wind tunnel. The SR-3C-3 model was mounted on an isolated axisymmetric nacelle test rig in the NASA-Lewis 2.44 x 1.83 m (8 x 6 ft) wind tunnel. This rig was capable of orienting the rotor drive through a range of tilt angles relative to the tunnel axis, to provide non-uniform inflow excitation to the rotor. The rotor drive is the same rig that has been used to test all the model Prop-Fans at NASA-Lewis and contains a 746kW (1000 shp) air turbine. the wind tunnel and Reference 3 discusses the detail.

Reference 6 discusses nacelle test rig in greater

2.3 Model Instrumentation

Foil strain gages mounted on the camber (suction) surfaces of selected blades were used to measure strain due to blade flexure. The strain gages were mounted by NASA-Lewis personnel, based on guidance provided by finite element analyses.

The strain gages were located at points along the blade mid-chord where the stresses associated with the first four modes were calculated to be high. Figure 2 shows the locations of the strain gages as they were applied to the blade. The gages were used to measure inboard bending, mid-blade bending, and mid-blade torsion (shear strain). Two opposing blades (numbers 1 and 5 ) had a full complement of the three gages. In addition, selected gages were installed on blades 2, 3 and 6, as described in Table 11.

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The strain gage signals were transmited from the rotor to the fixed frame system using a rotary transformer device. to magnetic tape recording equipment.

The output was ultimately directed

2.4 Test Procedures

Initially, the tunnel was brought up to speed with the Prop-Fan windmilling (zero power). pitch angle setting and tunnel free stream velocity. The model rotational speed, at this fixed blade angle and fixed tunnel Mach number, was incrementally increased by increasing the power to the rotor. This was done until an operating limit, such as a blade stress limit, rotational speed limit, or rig power limit was reached. angles, with the tilt angle being varied from the control room. process was repeated for different Mach numbers, also varied from the control room.

Prop-Fan windmilling rotational speed is dependent on the blade

This was repeated for various shaft tilt The whole

The tunnel was shut down in order to change blade pitch angle (ground adjustable). reference location on the blade (reference blade angle) prior to tunnel start up. The reference location for the SR-3C-3 is at the 0.78 radial station. The blade/hub collective pitch mechanical arrangement allowed for the adjustment of all blades simultaneously. average of the measurements for all of the blades.

An inclinometer was used to set the blade pitch angle at a

The blade angle was defined as the

2.5 Test Conditions

The conditions for these wind tunnel tests, after the tunnel has reached steady state operation at between 0.7 and 0.9 Mach number, include an air density equivalent to a standard day altitude of between 1524 to 2134 meters (5000 to 7000 feet). This is close to sea level conditions, as compared with the Prop-Fan design cruise operating condition at 10668 meters (35000 feet) altitude.

The parameters that were variable for the test were Mach number, Prop-Fan shaft tilt angle, blade angle and rotor RPM. All of these parameters were remotely controllable from the control room, except blade angle. A schedule of the Mach numbers, blade angles, and rotor shaft tilt angles which were tested is found in Table 111. The RPM's tested are also found in the table and they range from 3730 RPM to 10000 RPM. The RPM was tested in 500 RPM increments, from the windmilling RPM to the upper RPM limit. Figure 3 shows the operating envelopes for this test. These boundaries include the RPM limits encountered, determined by windmilling, the maximum drive power available or blade steady loading safety limits (9000 RPM). An exception to the 9000 RPM limit was made in order to probe for flutter at 0.9 Mach number, where 10000 RPM was allowed. The upper bounds on rig tilt angle and blade angle were generally limited by high strains. One set of boundaries is shown for each Mach number tested.

2.6 Data Reduction

Two types of magnetic data tapes were provided t o Hamilton Standard by NASA-Lewis. One contained the operating condition data found in Table I11 in digital form, and the other contained the strain data, in analog form, for all of the gages. The first type (condition data) was used during the data

4

reduction process to formulate the operating condition tables and data trend summary curves.

The second type (strain data) was also processed at Hamilton Standard using a computer based instrumentation data tape playback system. signals were passed through a scaling amplifier and then through peak detectors. specific time intervals, and the averaged half-amplitude was determined. peak detector output was sampled by an analog-to-digital converter, calibrated in engineering units and stored in computer memory. processed by a computer based analysis system.

The strain gage

Positive and negative peak strain amplitudes were averaged over The

The data were then

Once the sampled data resided in memory, a statistical treatment of the data was used to define the "total strain". For the present work, a level defined by the half-amplitude mean plus 2 times the standard deviation was used. That is;

€total = xbar + 2 * sigma. The instantaneous strain will be below this level 97.72 percent during the data sampling period. strains are above this value.

of the time That is, only 2.28 percent of the vibratory

The core of the data analysis system is a high speed mini-computer. computer was used to store the total strain data on a dual rigid disk drive. These data were later used to create trend summary plots of total strain VS. RPM and other test operating variables.

This

The data analysis system has the capability to perform spectral analyses of the analog signals. for every steady state run analyzed. at Hamilton Standard, identified the peaks, above a specified noise level, from the spectral data. multiples of the rotational speed) and trend summary plots were made from these data, and will be discussed later in the report.

The spectral data (in digital form) were stored on the disk An algorithm for the computer, developed

Tables of P-order strain values (strains at integer

3.0 ANALYTICAL TECHNIQUES

3.1 Approach

Extensive use was made of the MSC/NASTRAN finite element analysis computer program, described in reference 7, for the 1-P structural dynamic analysis of the model blade. Careful modeling techniques are required in order to create the finite element grids necessary to describe the Prop-Fan blade.

For most of the calculations, a finite element model for the SR-3C-3 blade, provided by NASA-Lewis, was used. and a schematic representaton of the model is shown in Figure 4. project, an improved finite element model was generated by Hamilton Standard using CQUAD4 elements. It is also shown in Figure 4. calculations were performed using this model. model was based is described in Reference 8.

This model is composed of CTRIA3 elements, Later in the

A limited number of The study on which the improved

The theoretically based methods used for this study, to predict the 1-P response of the blade to angular inflow, have been used in previous Prop-Fan structural dynamic studies, as described in references 2 and 3. discussion of these methods is given here.

A short

Figure 5 shows a block diagram of the prediction methods used in this analysis. The computer codes used in this analysis are listed in Table IV, where they are matched to their numerical designation.

Starting at the top of Figure 5, the model description, steady airloads (as computed by the HS/H039 and HS/H045 codes), and centrifugal load input into MSC/NASTRAN to determine a steady displaced blade position. airloads were computed for the angular inflow conditions, using the HS/H039 flow field analysis and the HS/H337 skewed wake analysis. distributed over the finite element model using HS/F194, and input into the MSC/NASTRAN structural dynamics analysis . determine the blade strain at the gage locations.

effects were The 1-P

These airloads were

A post-processor code was used to

3.2 Calculated Modes and Frequencies

Blade mode shapes and frequencies were determined by NASA-Lewis personnel using their NASTRAN CTRIA3 model for the SR-3C-3. calculated modes, in order of their respective frequencies are shown in Figure 6a, for the non-rotating (zero RPM), and the 8600 RPM conditions. frequency is shown beneath each modal pattern. for the non-rotating condition, and the first four modes are shown for the 8600 RPM condition. It is seen that there is little difference in the calculated mode shapes between the zero and 8600 RPM conditions, for the first four modes. The differences in the calculated frequencies are significant, however, showing the effects of centrifugal stiffening. Centrifugal stiffening raises the modal frequencies.

Schematic diagrams of the

The modal The first six modes are shown

This effect is greatest for the lower modes.

Also shown in Figure 6a are tracings of holographic patterns from photographs taken during a vibration test. This test was conducted at NASA/Lewis, on a non-rotating SR-3C-3 model blade vibrating at its natural frequencies. The measured frequencies are shown beneath each figure. Although the measured and

7

calculated mode shapes are generally similar, differences are seen for all six modes. displacement pattern compared to the holographic data. The fifth measured mode is similar to the sixth calculated mode, which appears to be a second torsional mode. The correlation between the calculated and measured non-rotating frequencies is good for all of the modes, except the third and fifth modes. Here, the frequencies differ by almost 15 percent. differences affect the values of calculated blade strain.

In particular, the fifth calculated mode shows a much diEferent

It is not known how these

Calculatons of blade mode shapes and frequencies for the zero RPM condition, made by Hamilton Standard using the CQUAD4 model, are shown in Figure 6b. The shapes derived using the CQUAD4 model are similar to those for the NASA CTRIA3 model. However, the CQUAD4 model frequencies generally match the holographic test data more closely.

The critical speeds are seen in Figure 7, which is a Campbell diagram showing the calculated frequency data, discussed above. Generally, these frequencies are typical of Prop-Fan models. observed during the wind tunnel tests, later in the report.

They will be compared to the rotating results,

3.3 Calculated Vibratory Strains

The results of vibratory response calculations for six cases are given in Table The analytical predictions were performed at the measured wind tunnel conditions. These test cases were selected so as to provide a range of operating conditions, such as RPM, Mach number and rotor power, so that important trends could be identified. number, and therefore the calculated cases are designated in the same manner. The operating condition for each case.

V. These cases were selected from the test points given in Table 111.

The test runs are listed by a reading

parameters of air density and temperature are shown

Blade strains were calculated using the NASA-supplied CTRIA3 NASTRAN model for all six selected conditions. The improved CQUAD4 model was used to calculate strains for case number 6 only. This case is most closely associated with the design operating condition.

The measured strains for each case which was selected for prediction are also listed in Table V. given in Section 4.8. response caused by the periodic aerodynamic loading excitation due to angular inflow. discussed. positions, inboard bending, mid-blade bending and mid-blade shear (torsion). The blade responses are given as values of strain divided by excitation factor (EF), a quantity which is sometimes known as The use of EF for normalizing strain data is intended to account for the dependence of strain on inflow angle and flight velocity (dynamic pressure). This was shown to be good practice in previous studies (References 2 and 3 ) , and will also be discussed later in this report.

A detailed comparison of measured and predicted strains is The calculated strains represent the vibratory blade

The strains are calculated at the strain gage locations, as previously These locations represent the following strain measurement

11 strain sensitivity".

Review of the calculated blade strain sensitivities, shown in Table V, reveals them to be only weakly dependent on changes in operating condition. sensitivity varies little over the range of RPM and Mach number studied. This is partially a result of normalizing the strain by EF.

Strain

The only significant

8

change is the increase in strain sensitivity seen in cases 1 and 6 at conditions of large shaft power. probably related to a change in the spanwise distribution, as well as the magnitude, of the aerodynamic loading on the blade. The trends of strain with operating condition will be examined further in the discussion of the measured test data, to follow.

The reason for this is not clear, but it is

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4.0 TEST DATA EVALUATION AM) COMPARISON WITH CALCULATIONS

The objectives of this SR-3C-3 Prop-Fan model test were:

1 ) To demonstrate the effectiveness of structural tailoring, using composites, in preventing blade instability (flutter clearance).

2 ) To determine the effect of pure angular inflow on the vibratory response of composite material Prop-Fan model blades.

To verify and evaluate theoretical calculations by comparison to test results.

3)

4 ) To compare the SR-3C-3 test results to the results of other Prop-Fan model tests.

4.1 Flutter Clearance

The SR-3C-3 Prop-Fan was tested, at zero tilt angle, at flight speeds up to 0.9 Mach number and rotational speeds up to 10000 RPM. be free of unstalled flutter instabilities over this entire operating range. This model was intended to be stable over a large portion of its operating range, by virtue of the structural tailoring of its composite ply layups. Further discussion of the stability analysis of this model is given in Reference 5.

It was demonstrated to

4.2 Total Strain Measurements

Blade strain measurements were made as described above, during wind tunnel testing on the Prop-Fan operating with its drive shaft tilted relative to the tunnel centerline, to provide angular inflow to the rotor. previously (Section 2.6), the measured total strain amplitude was extracted from the data using a statistical approach. half amplitude mean plus twice the standard deviation (xbar + 2 * sigma). Total vibratory strain measurements were obtained at steady state operating conditions. the gages, which are listed by reading number. data sample taken at a single operating condition. these runs represent are found in the performance table, Table 111, as discussed in Section 2.5.

As discussed

It represents the vibratory strain

Appendix I contains a table of the total strain values f o r all of A reading number identifies a

The operating conditions

For this study, plots of total vibratory strain were made, from data at all the steady state conditions. Total strain was plotted as a function of rotational speed (RPM) for various tilt angles, and combinations of blade angles and Mach number. Samples of these plots are shown in Figure 8, which contains plots of total vibratory strain as a function of rotational speed, at a Mach number of 0.7 and a blade angle of 59.0 degrees. 3.0, and 5.0 degrees are shown.

These data indicate that blade strain increases significantly with increased tilt angle, as expected. condition, probably due to several causes.

Data for tilt angles of 0.0,

There is some residual strain for the zero tilt angle There may be a small angular error

11

in the physical rotor shaft alignment to the tunnel centerline, and the pylon/nacelle may be introducing a small degree of non-unif ormity to the tunnel flow. Also, blade response may be induced by the small amount of tunnel turbulence present.

Note, in Figure 8, that a strain peak occurs at or near 8000 RPM. to the first mode/2-P critical speed crossover. critical speed is indicated in Figure 7 to be lower, at about 7000 RPM. measured critical speed will be discussed in the next section.

This is due The calculated first mode

The

4.3 Spectral Analysis

Spectral analysis of the strain gage signals is a useful tool for identifying the harmonic P-order and non-P-order (modal) responses of the blade. Spectral analyses were conducted for all of the steady state runs, the data for which were stored permanently on computer disk. For this study, spectral plots were made from these data for selected test runs, for gages BG1-1, BG1-3, BG5-2 and BG5-3. These data represent respectively, inboard bending and shear on blade number 1, and mid-blade bending and shear on blade number 5.

Figure 9 shows typical samples of the spectral plots, giving strain amplitude as a function of frequency for a rotational speed of 7300 RPM, a Mach number of 0.7 and a blade angle of 59.0 degrees. was 5.0 degrees. The plots show significant P-order response, while at the same time showing low amplitude modal response. Note that the shear strain spectra fo r the two blades are quite similar, showing the consistency of the test data.

The tilt angle for this run

The 1-P blade loads dominate, as was expected, since the aerodynamic loading is due primarily to pure amplitude, probably due to the fact that this condition is near the 2-P/first mode critical speed. time. sweep and flexibility. Further study of these data, and the test data which exist for other Prop-Fan models, would be helpful in clarifying the cause of this phenomena.

angular inflow. The 2-P loads are of significant

The source of the 2-P excitation is not known at this The 2-P response may be evidence of non-linear effects, due to blade

It is seen from the spectra in Figure 9, that non-p-order peaks are of very low magnitude. define the modal responses. to the small amount of turbulence in the wind tunnel, which is the source of random excitation for the blades.

These peaks, and their underlying broadband humps, were used to The low level seen for the modal responses is due

4.4 Modal Frequencies and Comparisons to Predictions

Blade modal frequencies for rotating operating conditions were identified using the blade strain signal spectra, described previously. Campbell diagrams illustrating the blade modal response were generated and are shown in Figure 10. speed, for data from the inboard bending, mid-blade.bending, and shear gages. Also shown on these plots are the calculated frequencies (see Section 3.3), and the holographically measured frequencies at zero RPM, as supplied by NASA-Lewis.

Blade modal response frequency is plotted as a function of rotational

Calculations of modal frequency, as a function of RPM, were

12

performed using the CTRIA3 NASTRAN model, holographic measurements were performed f o r the zero RPM condition.

The CQUAD4 calculations and

Inboard bending gage. bending gage clearly identiry the first mode. Data for the higher order modes, however, contain much scatter. first bending mode, with little response to the other modes, which accounts for the lack of clarity in the higher modes. Also, as was noted earlier, the modal responses indicated by the spectral data were of small magnitude, and thus difficult to identify, which could add to the data scatter.

The modal frequency test data indicated for the inboard

The inboard gage responds primarily to the

It was observed that the measured first mode is a stronger function of rotational speed than the CTRIA3 model calculated curve. the test response frequencies for the first mode were higher than predicted. This indicates FEA modelling deficiencies.

At the higher speeds,

Calculations of modal frequency as a function of RPM, using the improved CQUAD4 model, would be helpful in verifying the improvement in modelling centrifugal effects in the structural analysis. Also, the addition of aeroelastic effects in the structural analysis would improve the prediction of modal response frequencies.

Mid-blade bending gage. The modal responses indicated by the mid-blade bending gage are more consistent than the inboard gage data. located in a position on the blade that is predicted to respond more readily to higher mode excitation. the calculations show only four modes over the same frequency range. The cause of the third experimental mode is not clear. amplitude, as are the inboard gage data.

The mid-blade gage is

The test data indicate five response modes, whereas

These responses are of very low

Correlation between test and prediction indicates that the measured second, and higher order mode frequencies are substantially lower than the CTRIA3 calculated values, although they have similar slopes. Again, the measured first mode has a steeper slope and higher frequencies than the calculations.

Shear gage. mid-blade bending gage, except that it gives only weak indications of a mode in the 550 hz region. As before, the measured first mode response frequencies are higher than the calculated values. frequency values than the predictions. for improvement in the CTRIA3 structural model, in order to more accurately represent the response behavior of the blade.

The shear gage modal response is similar to that seen for the

The higher mode test data show lower These results again indicate the need

4.5 P-Order Vibratory Strains

A computer code, developed by Hamilton Standard, was used to search the spectral data stored on disk, to identify the strain peaks, and determine their values. plotting. study, the minimum value was 0.5 micro-strain. P-order values of vibratory strain, tabulated according to reading number, along with selected operating parameters.

These "peak values'' were also stored on disk for tabulating and The only peaks saved were above a minimum strain level. In this

Appendix I1 is a listing of the

13

For this study, with the Prop-Fan operating in an environment of pure angular inflow (except for nacelle and pylon effects), the dominant 1-P blade responses are of primary interest.

Effect of rotational speed. through 16. Here, for the three gages on blade number 5 .

The 1-P data have been summarized in Figures 11 micro-strain is plotted as a function of rotational speed

These figures display test data for operation from 0.36 through 0.85 Mach number. each Mach number condition are given in Table VI. show the 1-P vibratory strain increasing with rotational speed for the inboard bending gage and the mid-blade bending gage. amplitudes are consistently higher.

The combinations of Prop-Fan shaft tilt angle and blade angle used for The test data generally

The inboard gage strain

The shear gage data behave in a similar manner, up to 0.7 Mach number. point, the 1-P vibratory strains tend to decrease with increasing rotational speed. chordwise aerodynamic center location. loading, while not affecting the bending loads. although the mid-blade bending gage is located very close to the shear gage, it does not show this drop-off.

At that

This may be an effect of compressibility, due to a shift in the This would modify the torsional

It should be noted that,

Effect of Mach number. The effect of Mach number on 1-P strain data is shown in Figure 17. Here, 1-P micro-strain measured by the inboard bending gage is given-for the rotor operating at about 7000 RPM, at a tilt angle of 4 degrees. As indicated, some of the data were measured directly at 4 degrees tilt and some were interpolated from test data measured at other tilt angles.

The data show that 1-P micro-strain increases rapidly up to about 0.8 Mach number, and then levels off. The initial increase is due to dynamic pressure effects. than dynamic pressure at high speed. These stronger effects could include the decrease in angle of attack difference, between the advancing and retreating blades, with increased forward speed, and compressibility effects.

The leveling may be due to other effects, which become more important

Effect of blade angle. little effect on 1-P strain.

The test data show

4.6 Shaft Tilt and the Excitation Factor

The effect of ProD-Fan rotor shaft tilt is

that blade angle (rotor power) has

shown in Figures 18 through 20, where 1-P and 2-P-vibratory strain are plotted as functions of tilt angle. Each set of points has been fitted by a hyperbolic curve generated by a computer algorithm. The curves are nearly linear, however.

Figure 18 displays data for 0.6 Mach number operation at a blade angle of 57.0 degrees. The figure contains plots for test data at 5500, 6000, 6500 and 7000 RPM. The data shown in Figure 19 were taken during 0.7 Mach Number tests, for a blade angle of 59.0 degrees. number tests, for a blade angle of 61.0 degrees.

Figure 20 displays data taken during 0.8 Mach

The test data trends show increasing vibratory strain with increasing tilt angle. This trend is expected since increasing the tilt angle increases the

14

difference in angle of attack between the advancing and retreating blades. angle of attack difference is tlie source of the vibratory aerodynamic loading. There is a linear dependence on shaft tilt angle, away from the origin, €or both 1-P and 2-? strains.

The

Excitation Factor. suggests the use of the term "Excitation Factor" for the analysis of data trends.

The linear dependence of 1-P blade strain on shaft tilt

Excitaton Factor (EF) is defined as: 2

EF = ( Veq / 348 )

where airspeed in knots. pressure.

is the tilt angle in degrees and Veq is the equivalent sea level EF is proportional to Prop-Fan shaft tilt angle and dynamic

For a uniform, steady inflow to an untilted rotor, theoretically there is no aerodynamic excitation to induce a forced response of the blades. shaft is tilted at some angle to this uniform flow, a sinusoidal variation in velocity at the blade will occur with a frequency of 1-P. 1-P airload on the blade, that is some function of the mean flow velocity and density (dynamic pressure) and the shaft tilt angle.

If the rotor

This will produce a

For most operating conditions of interest, that is, away from critical speeds, it was shown above that blade stress is a linear function of shaft tilt angle. Past study has shown a linear dependence of blade stress on dynamic pressure, also. This allows the use of EF to normalize blade stress. In fact, this concept has been in use for many years. Other demonstrations of this concept for Prop-Fan data are given in References 2 and 3. Note that the quantity strain divided by EF is sometimes known as "strain sensitivity".

4.7 Power Coefficient

The effect of power variation on blade strain can be studied through the use of the term "power coefficient". application to propeller data analysis. non-dimensional function of the dynamic pressure, due to rotational speed at the blade tip, and diameter cubed. power the rotor absorbs is proportional to the tip dynamic pressure and diameter cubed. Power coefficient is defined as:

This term has been in use for many years, in The power coefficient is a

That is, everything else held constant, the

2 3 C p = 2 " f Q - - ?f3 Q

2 5 D 1/2 e Vtip D

3 e n

where p = air density in kg/m , Q = rotor torque in n-m, n = rotational speed in revolutions per second, Vtip = blade tip rotational speed in m / s , and D = rotor diameter in m. Use of the power coefficient normalizes tlie effect of rotor size and speed in the data. In the range of linear aerodynamics, the power coefficient includes the effect of blade angle.

15

4.8 Strain Sensitivity VS. Power Coefficient

A summary of the test data, normalized as described above, is given in Figure 21, where for each Mach number, the data for all blade angles, a11 RPM's and all tilt angles are shown in one plot. sensitivity plotted as a function of power coefficient, from the output of the inboard bending gage. forced excitation.

Figure 21 displays 1-P strain

This gage had the highest response amplitudes due to 1-P

These curves summarize the data for each Mach number tested. For Mach numbers of 0.6 or higher, the data collapse into a single curve. number condition, the effect of blade angle appears significant, with the data forming a different trend curve for each blade angle tested. Strain sensitivity decreases with increasing blade angle for constant power coefficient.

For the 0.36 Mach

In general, the curves fitted to the data are parabolic, with strain sensitivity increasing somewhat as the power coefficient increases. windmilling points (zero power coefficient) at 0.6 Mach number and above, the strain sensitivity value is. consistently about 60. sensitivity with power coefficient are remarkably similar over the Mach number range tested. number, up to 0.8 Mach, and then decreases slightly.

For the

The trends of strain

Strain sensitivity increases somewhat with increasing Mach

4.9 Comparison of 1-P Measurements to Predictions

Calculated strain sensitivities for six conditions are shown in Figure 21, in addition to the measured test data. These operating conditions correspond to those for six test runs (see Section 3.3) . cases using the CTRIA3 NASTRAN model. Calculations using the improved CQUAD4 model were made for the 0.8 Mach number (design point) case, only.

Calculations were made for all six

The blade angle used for the 0.36 Mach number case was 47.7 degrees. The calculated strain sensitivity for this case is considerably higher (88 percent) than the measured sensitivity, as shown in Table V.

For the higher Mach number conditions, the CTRIA3 calculations also significantly overpredicted the measured values (50 to 95 percent). improved CQUAD4 model reduced the overprediction to about 33 percent at the design point condition (case 6A).

The

Overprediction of measured blade response was not evident in studies of the dynamic response of metal Prop-Fan blades (Reference 2, 3 and 4 ) . causes may be responsible. Non-linear effects, evident from the significant 2-P responses, discussed earlier, are not included in the prediction methodology. effects. previous Prop-Fan model testing should be examined, to determined the extent of non-linear and aeroelastic effects. future improvements to the calculation procedure.

A number of

Also not included are twist magnification and other aeroelastic Although beyond the scope of the present study, data from this and

If important, these should be included in

16

4.20 Comparison to other Prop-Fan Models

Dynamic response test results for the SR-3C-3 were compared to results for other blade models, previously tested. This comparison was made with blade strain sensitivity data derived from other Prop-Fan tests conducted in the NASA-Lewis 8 X 6 wind tunnel, using the plotting format of Figure 21. given in Figure 22, which shows a set of plots, by Mach number, of blade strain sensitivity as a function of power coefficient. Comparisons are shown with the SR-2C composite material model, and the SR-3 and SR-5 solid titanium models, all of which were discussed in Reference 2.

This is

The strain sensitivity of the swept composite SR-3C-3 model is generally lower than that of the straight composite SR-2C model. This indicates the benefit of blade sweep in reducing blade response. The benefit may disappear at extremely high blade sweep, noting that the highly swept SR-5 model strain sensitivity is generally higher than that for the moderately swept SR-3 model.

As shown in Figure 22, the composite material SR-2C and SR-3C-3 blades have greater strain sensitivities than the solid metal SR-3 and SR-5 blades. may be due, in part, t o diff.erences in blade material stiffness and.inertia properties. In addition, the non-linear effects discussed earlier may influence blade responses.

This

Further studies are needed to define this behavior.

5.0 CONCLUSIONS

As a result of this study of SR-3C-3 model Prop-Fan blade dynamic response and stability, the following conclusions are made:

I 2,

7)

This composite blade was structurally adequate over its entire operating range.

The swept composite blade showed less response than the straight composite blade.

The trends of l-P blade response were defined using non-dimensional data (strain sensitivity VS. power coefficient).

Composite blades have higher strain sensitivity than metal blades.

l-P blade response was overpredicted 50 percent at the design point, and up to 95 percent at off-design conditions, using unimproved methods.

An improved finite element modelling method reduced overprediction of l-P response to about 33 percent, at the design point operating condition.

High 2-P response levels suggest the existence of non-linear effects, not included in the prediction methodology.

19/20

6.0 RECOMMENDATIONS

I Based on the conclusions of this study, the following recommendations are made:

The existing test data for all Prop-Fan models should be reviewed to determine the non-linear effects on blade response.

Non-linear effects should be included in the blade response prediction methodology.

The improved CQUAD4 FEA model should be used for additional calculations of SR-3C-3 blade modal and forced response.

2 1/22

7.0 REFERENCES

1.

,

2.

~

I

I 3.

I I

I 4. I

I

I I

5.

6.

7.

8.

Gatzen, B.S. ; "Prop-Fan Potential and Technology Summary", Presented at the Japan Society for Aeronautical and Space Sciences meeting, Tokyo, Japan, November, 1984.

Bansal, P.N.; Arseneaux, P.J.; Smith, A.F.; Turnberg, J.E. and Brooks, B.M.; " Analysis and Test Evaluation of the Dynamic Response and Stability of Three Advanced Turboprop Models", NASA CR-174814, August, 1985 . 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.

Smith, A.F.; "Analysis and Test Evaluation of the Dynamic Stability of Three Advanced Turboprop Models at Zero Forward Speed", NASA CR-175025, December, 1985.

Turnberg, J.E. ; "Unstalled Flutter Stability Predictions and Comparisons to Test Data for the SR-3C-X2 Model Prop-Fan", NASA CR-179512, October, 1986.

Swallow, R.J. and Aiello, R.A.; "NASA Lewis 8 X 6 Supersonic Wind Tunnel", NASA TM X-71542, May, 1974.

MacNeal, R.H.; " The NASTRAN Theoretical Manual", (Level 15.5), MacNeal Schwendler Corp., December 1972.

Smith, A.F. and Brooks, B.M.; "Dynamic Response of Two Composite Prop-Fan Models on a Nacelle/Wing/Fuselage Half Model", NASA CR-179589, October, 1986.

23/24

TABLE I

DESIGN CHARACTERISTICS OF THE SR-3C-3 MODEL PROP-FAN

Parameter Value

Diameter, cm (in) 62.2 (24.5)

Activity Factor/blade, AF 235 Activity Factor, Total 1880 Airfoil series (NACA) outboard 16

inboard 65/CA Integrated design lift coefficient, C1 0.214

Material Carbon/Epoxy fiber composite

Number of blades a

Blade aerodynamic tip sweep, degrees 34.5

Fiber orientation (l), degrees 0, +45

Cruise Conditions:

Speed, Mach number Altitude, km (ft) Power loading, kW/m2 (shp/f t2) Tip rotational speed, m/s (fps) Power coefficient, Cp Advance Ratio, J Cruise efficiency, percent Cruise noise ( 2 ) , dB

0.8 10.7 (35,000)

300 (37.5) 244 (800)

1.695 3.056 78.4 144.5

(1)

(2)

Zero degrees fiber orientation parallel to pitch change axis.

Maximum sideline noise at blade passage frequency.

TABLE I1

STRAIN G A G E DESIGNATIONS FOR SR-3C-3 RESPONSE TESTS

R a d i a l G a g e s t a t i o n B l a d e n u m b e r D e s c r i p t i o n cm 1 2 3 4 5 6 7 8

I n b o a r d B e n d i n g 11.9 B G 1 - 1 BG2-1 B G 3 - 1 - B G 5 - 1 B G 6 - 1 - -

M i d - b l a d e B e n d i n g 2 4 . 6 BG1-2 BG2-2 - - BG5-2 BG6-2 - -

S h e a r 2 6 . 0 B G 1 - 3 - - - B G 5 - 3 - - -

G a g e s a r e d e s i g n a t e d BGx-y, w h e r e x = b l a d e n u m b e r a n d y = g a g e n u m b e r .

H H H

w ell m 4 H

w ell 3 a w 2: u cn

u f+ c13 w h cr, 0

VI I- W w c

I c-l L w Z m 4.2- c 3 3 C C W t - l GI

m z < I l-4W

L? 3 c

rr.0 x

m

a

7.

U I 0 1 0 1

I .................................................. n ~ m ~ o r l ~ m m ~ ~ a ~ ~ ~ ~ . h h ~ h ~ ~ ~ o . m m m o r ~ a ~ . ~ m a m ~ ~ m m o ~ - ~ o ~ ~ ~ m m a u o . -

L w r I h N rl N o 4 rl c o N h 0.m N N P o r d rd o - .-t o 4 0 o Li) h o o w L I o o - o m o m o m - Q Q o -a Q o n o a o M o m o m m o y1 o m 01 o h o m o n e ~o c o c E n o o m r d c+ o EL E I m 01 M P a h Q c m P P 01 P c D. a or\ h Q -a m n P P m m Q Q m o m o o *m m Q Q Q P Q I, Q bo. k~ :? m 3.

N D. r d N O N o -a c m o r. - o o - - N v

P W

I h o - o o 0.- - r( r.1 e rdm - r d - (u ri N - r a - m r q ri r.i P m m m o. o. - r d - r d N r( r.1 a a a o. x hh m m m o I '1'1 'I q '1'1 'I 'I 'I q '1 'I 'I? 'I'I 'I ,??'I 'I 'I '1 'I '1 'I'I 'I'I q '1 ?'I '1; ?'I 'I m n ll? n b? "7 bY b7 < u .Q m I

I

M Q u u Q L-7 Q Q u .c Q Q .a .Q Q "3 Q "3 .o Q Q Q Q Q .a u Q u Q Q M n Q Q u Q Q Q u m 0 a m 0 m 0.0 0 0 0.

1 o o o o o o o c o o o o c o o o o o o o o o o o o o o o o o o o o o c o o o o o o o o o o o o c o o . . . . . . . . . . .

27

La c La W c

2 C A I L W

zu1 O Z U

K l - W L J

C I m z c I Y V I v a c I 9 2 I - K - 4 La

X

0 3 3

m

I I

I - U O C Z x

a

e 2

Y Lc

28

vl I- u) W I-

2 C - l L r W zua 0 2 - C 3 X K I - W L A G I m z c I - m O S <

K O X vl

7 .

m

I I h o. o P n -a -a P J m N -r.i n ri v) lil Q E o o r d N 4 r d - r.1 r d r d o P E rl o M o P o @ l i ~ E - o r( o o o -a s

u o < z I ~ ~ ~ ~ ~ ~ : I C - ~ ~ I C ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ Q ~ ~ Q Q O ~ Q Q ~ Q ~ ~ ~ ~ ~ ~ ~ ~ ~ G I G I I . . . . . . . . . . . . . . . .................... z - 1 o . P P P P o . P P o o o ~ o o o o . o c ~ E o o o o c o o c o o o P P ~ Y ~ P P P c 5 o . P P E P ~ P P P i ?

s I c o o c o o o o o o o o o o o o o o o o o o o o o c o o o o o o c o o o o o o o o c o c c o c o o c

29

H H H

W

m 4 h

30

I in ................................ P I - N ~ . - ~ O ~ . L O * ~ ~ ~ O ~ L O ~ ~ - C - ~ - O ~ ~ O L O * ~ * N ~

< a C W L I o n o . ~ ~ ~ n ~ o o . ~ ~ o o o o ~ o . n ~ ~ ~ o ~ o o o . o n ~ ~ ~ r i ~ h LL w D. a h n .Q .o .on h r) ~ i ) o. a m m YI o. LO

z m I L ~ I 0 zr I

W 2 + L W s I O O 0 . O O O C ~ 0 . C n ~ 0 ~ 0 0 o . o C o m 0 . o 0 . m o o m o ~ m o

K L LT I o. a n n n .o n m o P i .a o.

m

I of ............................... c o w ~ m a m m o m o a m a ~ m m a ~ a m m m m ~ m a m m o ~ ~ ~ ~ ~ r ) -I z ci I o o o o o P P P P - o n LO nu> Lo n Y) b i ii) Li i n n LO II-J M -a Q -a -a a < I

1

- - G O c z S

0 z

s ic

o t m o coo. O E o m o -a o 7 b3 m 0.0 r.8 Q o. 01 o u 9 G U-J ri L~Y m n ~ i ) 21 -a -a . o b i h i n Q u o o. o e. c- o. E O o o o. o o. o E o. v -7 ~ i l rn o. ,m PY .. m q ~ m q q q q q I\ 2 p: 2 p .Q . . .............. -ab I\ I\ .o .a .o PI u Q I! a m Q m 6000060000Cd00000~00000000000000 %

* *

I

31

TABLE IV

HAMILTON STANDARD COMPUTER CODES USED FOR

BLADE DYNAMIC RESPONSE ANALYSIS

Code Designation Description

HS/H039

HS/H045

HS/H337

Potential flow field analysis, used to determine the influence of the nacelle on the inflow to the rotor.

Lifting line, quasi-static performance strip analysis, 2-D airfoil section data, Goldstein wake induction, azimuthal variations.

Lifting line, quasi-static performance strip analysis, 2-D airfoil section data, skewed wake induction, azimuthal variations.

HS/F194 Distributes airloads over finite element grid.

MSC/NASTRAN Finite element analysis used for calculating vibratory mode shapes and frequencies, and dynamic responses o f Prop-Fan model blades.

STRAINNP Converts element stresses from MSC/NASTRAN to strains at the strain gage locations.

32

z d u

E m I-

v) E 0 -r- L G O coo

L o u w -- m

v u X

w I 0 d

03 m

0 03 0

a3 QI

h d

d

cu h

d

m 0

II

d 4 0 W

v) z

v) W cr

2 L

W Y

U E m v) &

0 * 0 0 h

Io d a3

l n d

W I I-

E a,

5 LL 0 cu

h d

c3 h d

d cu d v,

I- 3 v) W e

d

U P a 0 a = L no,

m .-)

0 h

m 0 d h

d QI d W

d d 0 QI

d

> W

I- V E d cu

0 h

rD m

a v) m V

v) cu a3

m 0

h

QI 0

d

cu 0 v)

0 0

v)

W 0

c c c

cu cu

TABLE V I

SR-3C-3 MODEL 1-P VIBRATORY STRAIN TEST CONDITIONS

M e a s u r e m e n t s o f 1 - P v i b r a t o r y b l a d e s t r a i n d a t a , a t e a c h Mach n u m b e r t e s t e d , were a c q u i r e d a t t h e f o l l o w i n g c o m b i n a t i o n s o f s h a f t t i l t a n g l e a n d b l a d e a n g l e ( 9 3 / 4 ) , r a n g e . T h e d a t a a r e p l o t t e d a s 1-P m i c r o - s t r a i n VS. RPM i n t h e i n d i c a t e d f i g u r e .

o v e r t h e a l l o w a b l e RPM

T i l t B l a d e Mach A n g l e A n g l e

F i g u r e n u m b e r Deg . D e g .

11 0 .36 8.0 44.5 8.0 48.0

15.0 48.0 15.0 51.6

1 2 0.6 7 .0 4 .0 7 . 0 7 .0 7 . 2

54.6 57.8 57 .8 59.1 59.4

1 3 0.7 5.0 5.0 3 .0 5.0 5.0

57.8 58.4 59 .1 59 .1 61.5

14 0 .8

1 5

16

0.84

0.85

34

4.0 2.0 4 .0 4.0

2.0 4.0

4.0 4.0

59.1 61 .5 61.5 62.5

62.5 62.5

62.5 63.7

FIGURE 1. SR-3C-3 MODEL PROP-FAN INSTALLED IN THE NASA-LEWIS 8x6 TUNNEL.

35

BLADE POSITION 1

5

( v I EW I NG u PSTR €AM 1

26.0 cm I

24. b

Gaqe-gescEbiEti=2 -1 Inboard Bending -2 Mid-blade Bending -3 Shear

cm Camber Side

11.9 cm

Blade Pitch Axin

1 t t - \-& Prop-Fan Centerline

Figure 2 SR-3C-3 Prop-Fan Schematic showing the strain gage locations. NASA/Lewis 8 X 6 Wind Tunnel tests.

36

81 ade Angle deg.

50

40

B1 ade A n 9 1 e

deg.

\ P w --- -- - -O -

I 1 1 I I I 1 1

B1 ade Angle

deg.

60

50

B1 ade Angle

deg . -

I I I I I I I I

E l ade Angle

deg. 60

50

81 ade Angle

deg.

‘0- - - - - - - I_-

\

0,2,4 W

I I I I I

-

I 5

0,497

0,294

Figure 3 Test envelopes f o r t h e SR-3C-3 model Prop-Fan 8 X b wind tunnel tests a t N A S A / L e w i s .

= Strain Limit a t Max Inflow Angle W = Windmill Speed P = Rig Power Limit

37

n W

L H

M I 0 m

I cc v)

U

cc 0

1P DYNAMIC ANALYSIS

OPERATING CONDITION INFLOW ANGLE

I I AIRLOADS I AIRLOADS1 I I

HS/F194 I I I I DISTRIBUTE I I LOADS TO I I FE NODES I I I - - - - - - - - -

i IP 1 STEADY I

I I I I- I

I *

NONLINEAR STATIC ANALYSIS

SOLUTION 68 -----------------

FREQUENCY RESPONSE ANALYSIS

I I i

I I I POST PROCESSOR I

~~ I I

I STRAIN I I STRESS I I I

I I

I END I

I DEFLECTIONS I

FIGURE 5. DYNAMIC RESPONSE PREDICTION METHOD

39

I 31

I

I Ir'l I

I

I *I

I

N S r4 m d d

d

t4 I: 0. M m

N I: 0 d m

I W I

01 Zl

I I

ni I

MI I

t.1 N .c 9 0. 9

N rh b 9

I 1uI

I N r M

I 4

I k 9 N c

b N L M 0.

0.

0 u c CT C -0

c 3 3 0

Q J 4 J u l n a f 4 J

nul m Rj c r 1 1 0 m01E o r I l - l r c o a o I L Z

a c r d L 0 3 0112: u + r i 4 r c 6 9 0 O L L Z

Q VI .C

40

N I 0

00 7

N S

h 0 W

N I m d 4

N r

2=

3 3 5 s E a J

3 o w M a J

L - c e U a a J M U m o 3 E uc, l a w - -

Y I d h h

N I 0 m

N I cu cv d

41

X r.lSC/NASTRAN C A L C U L A T I O N S ( C Q U A D 4 )

-MSC/NASTHAN Calculations ( C T R I A 3 ) 0 Holographic Tests -NASA/Lewi s

Response Frequency

HZ

Propeller Speed - RPM

Figure 7 SR-3C-3 model Prop-Fan blade, natural f requenCie5

42

600 Micro Strain

400

200

- - 8

L.

-. - .II E l h ~ A

- r A A d - 1 1 I I I t I I I IJ

Figure 8

600 Micro Strain

400

200

Total Vibratory s t ra in vs propeller speed for the SR-3C-3 model Prop-Fan blade response t e s t s , NASA/ Lewis. Mach Number = 0.7, Blade Angle = 59.0 Deg.

- , - - - g o o

-.. A~~ E l

A A A * - El

- t I I I I L I r ?

43

Ln I

4

- Li3

02 0 91

'J

r I 1. ...............

r.4 I

Ir'J ................. m D C U c ................. 111

4

m 1

U L .d+J . Ern

- L17 .

.I

ul +J

ul 111 +J

m

111 L 3 D

LL .rl

44

h n -4- a M a 5 N a a

0 3 o r 0 a a l u 9

I C c a l 0 7 N

m a l alI n OI

a k

3 0 0 0 M

r O I L 0Is: 0 m i

'0 0111 OPI

911)

I OPI 0- O d * a l

n 0

3 L on. 0 c.(

o n

0 0

0 0 3 2 z

On.

0 a t o a

a U

a N

3 -

0

0 0 0 0 mal u 7 N o r $ 0 0

I C c a l 0 I 3 N

m o l n BI

alI a k

3 3 0 0 U

r O I L

3 0 0 1

'0 0 a l 3 9 911)

I 0al 0-4 0- t a l

3 L 0 a 3

o a

o n

E t-4

0

45

ONCE PER REV. 800.0

700.0

600.0

2 500.0 a a * 100.0

a r 300.0

H

0

u H

200.0

100.0

0.0

VIBRATORY S T R A I N

WCH. NO - ,36 BLRDE ANGLE ( . 7 5 R ) - 99.5 DEG, T I L T ANGLE 8 DEG,

a 0 O A A A

*. . . - . . ... I . . . - - - - . - 1 . - - . . - . - . I -..-. - - - . I . - - - . - - . - I - - - - - . - - ..rrrrrmrl

0 BGS-1 IN60 BEND 8 BGS-2 H I D BEND A BGS-3 SHEAR

0 II 0

A

890.6

700.0

600.0

IWCH. N(I - - 8 6 &&E AffiLE7(;75R) - 98.0 DEC. T I L T ANGLE - 8 DEC,

0 BG5-1 INBO BEN0 Q BGS-2 M I D BEND A BCS-3 SHEAR

z 500.0 a a * '100.6

H

+ 0 Q: u H s 300.0

200.0

890.6

700.0

600.0

z 500.0 a a * '100.6

H a s 300.0

H

+ 0 Q: u

0 m o a 0 o e b 3 % 200.0 o ! :

A A

100.0 A A

A A A

0.0 I . . . . . . - - . ................................................. - .

t A I A

0

e A A

e

A

0

b a 0 e e

a

% 0

A

0.3

-

0.Y 0.5 0.6 0.7 0.8 PROPELLER SPEED-RPM *lo'

0.9 1.0

F , y r e S R - 3 C - 3 BLADE DYNAMIC RESPON E TO ANGULAR INFLOW - 0.36 MACH

46

ONCE PER REV. VIBRRTORY STRAIN

800.0

700.0

600.0

z 500.0 a a v, ¶OO.O

a I 300.0

n

c

0

2

200.0

100.0

0.0

WCH. NO - .36 BLADE ANGLE (,75R) - '18.0 DEG, T I L T ANGLE - 15 DEG,

A

'"" " '. ~ , ~ . ~ . ~ ~ ~ . ' ~ . . . . . . . . . , . . ~ . ~ ~ ~ ~ ~ , . ~ ~ ~ ~ ~ ~ ~ ~ , ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ,

BG5-1 INBD BEND Q BG5-2 M I D BEND A BG5-3 SHEAR

m

MCH. NO - .36 BLADE ANGLE ( e 7 5 1 2 1 51.6 DEG.

700.0 T I L T ANGLE 15 DEG.

0 BGS-1 INBD BENO BG5-2 M I D BEND

A 8 6 5 - 3 SHEAR

600 0

z 500.0 a I4

z in 'Loo.0 0

u n a

r: 300.0

200.0

LOO. 0

A

a 8

e e e m

0 O b e

A A

A

A

Figure 11 SR-3C-3 BLADE DYNAMIC RESPONSE TO ANGULAR INFLOW - 0.36 MACH

47

' A

A A

800.0

700.0

600.0

z 500.0 a a - '100.0

a

x 300.0

H

I-

0

0 H

200.0

100.0

0.0

0 865-1 INBD,BEND Q 865-2 MID BEND A 865-3 SHERR

VIBRFITORY S T R A I N

mCH. NO - ,6 BLADE ANGLE (.75R) - 59 .6 DEG. T I L T ANGLE - 7 DEG.

. - - - . - - - - . - - - - - . -~.-- . - - . - -mrFh-.--- . . . . - . - . . - . . . - . . . . . . I

0 . 3 0.11 0.5 0.6 0.7 0.8 0.9 1.0

800.0

WCH, NO $ 6 BLADE ANGLE ( e 7 9 1 57 .8 DEG.

700.0 T I L T ANCLE 'I DEG.

600.0

2 500.0 H a a c - roo.0 0 a 0 H x 300.0

200.0

100.0 4

A

0.0 + . . . . . , . . . . . . . . . , . . . . . . . - . , . . . . . . . . - I . . . . . . " . I " ' . . . . . . i

0.3 0.q 0.s 0.6 0.7 0.8 0.9 1.0

PROPELLER SPEED-RPM *lo1

Q BGS-1 INBD BEND Q BG5-2 MID BEND

BGS-3 SHERR

Figure 12 SR-3C-3 BLADE DYNAMIC RESPONSE T O ANGULAR INFLOW - 0 .6 MACH

48

ONCE PER REV. VIBRATORY STRRIN

800.0

700.0

600.0

2 500.0 a a H

+ 0

0 H

900.0

a 300.0

200.0

100.0

0.0

MCH. NO - .6

, . - . - - - - - '.""...--.-----.-....,.........,...*.....,....~....,.~....~.~,

BLADE ANGLE-(,75R) - 57.8 TILT ANGLE - 7 DEG.

800.0

700.0

100.0

2 500.0 a

cn YOO.0

a

H

2 0

0 H 300.0

200.0

100.0

0.0

a BGS-1 IN80 EENO Q BG5-2 M I D BEND A BGS-3 SHEAR

WCH. NO - ,6 . a BGS-1 INBO BENO BLADE ANGLE ( . t S R ) - 59.1 Q BG5-2 MID BEND TILT ANGLE - 7 DEC. A 865-3 SHERR

n n l

A

A

8 s 0 0

A

A

F i g u r e 12 S R - 3 C - 3 BLADE DYNAMIC R E S P O N S E TO ANGULAR INFLOW - 0.6 MACH

49

m

800.0

700.0

600.0

z 500.0 a a n

b-

0

0 H

~ 0 0 . 0

a r 300.0

200.0

LOO. 0

0.0

m

A e

P

WCH. NO - .6 BLADE ANGLE (.75R) - S9.q T I L T ANGLE 9 7.25 DEG,

, . . . . . . . . . , . . . . . . . . . , . . . . . . . . . I . . . . . . . . . I - . . . . . . . . I . . . . . . . . . I . - . . . * . . - I

A

0 865-1 INBD BEND 0 865-2 MID BEND * BCS-3 SHEAR

Figure 12 SR-3C-3 BLADE DYNAMIC RESPONSE TO ANGULAR INFLOW - 0.6 MACH

800.0

700.0

600.0

2 500.0 a a H

I-

0

u n

900.0

a 300.0

200.0

100.0

0.0 0.3

000.0

700.0

600.0

500.0 a

v, 900.0

a

H

U b-

0

0 H 300.0

200.0

100.0

ONCE PER REV. VIBRATORY STRAIN

MCH. NO - . 7 BCAOE ANGLE (.75R) - 57.8 T I L T ANGLE - 5 DEG.

Q BG5-1 IN80 BEND Q 865-2 MID BEND * BG5-3 SHERR

e e e e

A A

mm ..............................................................., 0.11 0.5 0.6 0.7

PROPELLER SPEED-RPtl

MCH. No - .I BLADE ANGLE (.75) - 58.Y M G . T ILT fWGLE - 5 DEG.

0.8 0.9 1.0 *lo'

0

A

0

m e

e

A A

.................................................................. 0.3 0. Y 0.5 0.6 0.7 0.8 0.9 1.0

PROPELLER SPEED-RPH *lo'

F igu re 13 . SR-3C-3 BLADE DYNAMIC RESPONSE TO ANGULAR INFLOW - 0.7 MACH

51

BG5-1 INBD BEND Q 865-2 MID BEND A 865-3 SHEFlR

800.0

700.0

600.0

z 500.0 a a - '100.0

a

r 300.0

H

I-

0

u n

200.0

100.0

0.0 0.3

ONCE PER REV. VIBRATORY S T R A I N

a

6 e o m

e = a e o *

e * P A A A A

A

-1

0.11 0.5 0.6 0.7 0.9 1.0

WCH. No - . 7 ELROE ANGLE (.75R) - 59.1 DEC. T I L T ANGLE - 3 DEG.

800.0

mcH. NO - .7 700.0

600.0

z 5w.o a a H

I- - WO.0 0

0 H a

300.0

BLADE ANGLE-- 59.1 DEG. T I L T M L E - 5 DEG.

a 8GS-1 INBD BEND Q 865-2 M I D BEND A 865-3 SHEAR

0.3 0.Y 0.5 0.6 0.7 0.8 0.9 1.0 PROPELLER SPEED-RPtl *lo'

A A A A

200.0

100.0

0.0 0.3 0.Y 0.5 0.6 0.7 0.8 0.9 1.0

PROPELLER SPEED-RPtl *lo'

F i g u r e 13 SR-3C-3 BLADE DYNAMIC RESPONSE TO ANGULAR INFLOW - 0.7 MACH

52

800.0

ONCE PER REV. VIBRRTORY STRAIN

700.0 tmCH. Ma - .7 BLlDE ANGLE (.7SR) - 61.5 T I L T ANGLE - 5 DEG,

600 0

Q BC5-1 INBO BEN0 Q 8G5-2 M I D BEND A BGS-3 SHEAR

z 500.0 a a u-) 100.0

H

0

0 n a

s 300.0

200 0

100.0

b 0

0 t 4

0.6 . 0.3 0.q 0.5 0.6 0.7 i . 8 0.9 1.0

PROPELLER SPEED-RPPl *10

Figure 13 SR-3C-3 BLADE DYNAMIC RESPONSE TO ANGULAR INFLOW - 0.7 MACH

,

53

0 BG5-1 INBD BEND Q BG5-2 MID BEND A BG5-3 SHERR

800.0

700.0

600.0

2 500.0

a * '100.0

a

H a I-

0

u H 300.0

200.0

LOO. 0

0.0

I

WCH. NO - .8 BLADE ANGLE (e75R) 0 59.1 DEG. T I L T ANGLE 0 'I DEG.

, - - . - . - - - - . - - . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

800.0

700.0

600.0

2 500.6 H a U I-

0

u n

* 900.0

a

x 300.0

200.0

too. 0

0.0

m rn

MCH, NU - .8 BLADE ANGLE (.7SR) 61.5 TILT ANGLE 2 DEG.

a

e m

u r n a m e e e O . e r

A A . & A A

I . - - - - - - . - . ....,--..........-...-....-.......,-......'-...-...-'

0

A A

A

A

0 BGS-1 INED BENO Q BG5-2 MID BEND A BG5-3 SHERR

Figure 14 SR-3C-3 BLADE DYNAMIC RESPONSE TO ANGULAR INFLOW - 0.8 FlACH

54

800.0

700.0

600.0

z 500.0 a - '100.0

H Q: I-

0 K z 300.0

200.0

100.0

ONCE PER REV. VIBRATORY S T R A I N

WCH. NO - .8 BLADE ANGLE (.75R) - 61.5 DEG. TILT ANGLE - t DEG.

0

m a a

e . e e

A A

A

A

0 BG5-! IN60 BEN0 Q BG5-2 M I D BEND A BC5-3 SHERR

0.0 - .................................................. ---g

0.3 0.q 0.5 0.6 0.7 i . 6 0.9 1.0 PROPELLER SPEED-RPtl *10

800.0

700.0

600.0

z 500.0 a a - 100.0

H

I-

0

0 n a

300.0

200.0

100.0

MACH. NO - .79 BLADE AllGLE t . 7 9 ) - 62.5 TILT ANGLE 0 9 DEC.

e e

A A A

A

Q BGS-1 IN80 BENO Q BG5-2 MID BEND A BGS-3 SHEAR

h. .............. - 3 3 , 0.0 0 . 3 0.q 0.5 0.6 0.7 0.8 0.9 1.0

PROPELLER SPEED-APH *lo'

Figure 14 SR-3C-3 BLADE DYNAMIC RESPONSE TO ANGULAR INFLOW - 0.8 MACH

8 BGS-1 IN80 BEND Q BG5-2 MID BEND A BC5-3 SHEAR I

800.0

700.0

600 0

2 500.0 a a u) 100.0

a

H

c

0

0 H 300.0

200.0

100.0

0.0

mCH. No - ,811 BLADE ANGLE ( .75R) - 62.5 T ILT ANGLE - 2 DEG.

I n

~~0 g e l e

a 8 0 .

A A A ' A A

.........................................-........................ I

0.3 0.1 0.5 0.6 0.7 0.8 0.9 1.0

r. A

800.0

700.0

600.0

z 500.0 a a cn YOO.0

a

H

+- 0

u H 300.0

200.0

100.0

0.0

BC5-1 INBD BEND I

6 865-2 MID BEND A BC5-3 SHEAR

MACH. NO - ,811 BLADE ANGLE (.75R) - 62.5 T I L T ANGLE - 11 DEG.

......... I .........,......... I . . . . . . . . . , . . . . . . . . . , . . . . . . . . . , ......... 1

0.3 0.Y 0.s 0.6 0.7 0.8 0.9 1.0

E I

A

F i g u r e 15 SR-3C-3 BLADE DYNAMIC RESPONSE TO ANGULAR INFLOW - 0.84 MACH

56

800.0

700.0

600 0

z 500.0 a a cn ~ 0 0 . 0

a

H

+ 0

u H 300.0

200.0

100.0

0.0

a BG5-1 IN80 BENO Q BGS-2 M I D BEND A BG5-3 SHEAR

VIBRRTORY STRAIN

mcH. NO - ,as BLADE ANGLE (.75R) - 62.5 T I L T ANGLE - 't DEC.

..........,...................,-................,......... .

6 0

800.0

700.0

600.0

soo.0 H

+ 0

0

a a cn 100.0

a

300.0

200.0

LOO. 0

0.0

0 a

0 e

MCH. No - ,a5 0 BGS-1 INBD BENO BLAOE ANGLE t . 7 9 ) - 63.7 Q 865-2 MID BEND T I L T AWGLE - '1 MG, 4 BG5-3 SHEAR

..............................-,

e A A

A

e A

A A A

Figure 16 SR-3C-3 BLADE DYNAMIC RESPONSE TO ANGULAR INFLOW - 0.85 MACH

e Y

i w

a a n n

!-

w =I v,

H

X

;5

0 u)

m0 a D b

I I I I I I I I

I

0 0 d

0 2

m 0

- .

co 0

-.

h

0 - .

W 0

- .

u)

0 -

d

0 -

m 0

cu 0

-

-

7 _ . 0

. .

s + v,

z 0 I- s m Y . . > CL I 4

m I O m

I cc v)

z 0

I O s LL 0 I- O W LL LL w

h 7

Q) L 3 m

NIWl l IS -Ol j3 IW d-1

5 8

MODEL PROP-FRN TESTS METER WIND TUNNEL

NRSR LEWIS

1200.0 -

1 0 0 0 . 0 -

1000.0 -

800.0 - II \ II 0

I

0 K u U 15 600.0

l/REV. STRRIN Q Z/REV. STRAIN

z U

K t- LD

a

900.0

200.0

0.0

4

INBOARD BENDING STRAIN EC5-1

flACH NO -0.6 BLROE ANGLE - 57 DEG. PROPELLER SPEED - 5500 RPM 0 1/REV. STRAIN

8 2/REV. STRAIN

0.0 1.0 2.0 3 . 0 9.0 5 .0 6 . 0 7.0 PROPELLER T I L T FlNCLE - DEC

FIGURE lb .THE EFFECT OF PROPELLER T I L T ON THE 1P AND 2P VIBRATORY STRAIN

59

MODEL PROP-FRN TESTS 1 . 8 METER UINO TUNNEL

NRSR LEWIS w INBOARD BENDING STRAIN BG5-1

1 1200.0

1 1000.0

R A W NO -a6 BLADE ANGLE - 57 DEG. PROPELLER SPEEEO - 6500 RPfl 0 l/REV. STRAIN b, 2/REV. STRAIN

0 V H

a

600.0

2 H a a I- cn qoo.0

200.0

0.0

J

Q

0.0 1.0 2.0 3.0 11.0 5.0 6 . 0 7.0 PROPELLER T I L T RNGLE - OEG

HACH NO - a 6 BLADE ANGLE - 57 DEG. PROPELLER SPEED - 7000 RPH

0 l/REV. STRAIN Q Z/REV. STRAIN 1 1000.0

800.0 r \

0 V H

a

= 600.0

z H a a I- Lo 400.0

200 0 0

0.0

J

1

0.0 1.0 2.0 3 . 0 Y.0 5.0 6 . 0 7 . 0 PROPELLER T I L T RNCLE - DEG

FIGURE 18. THE EFFECT OF PROPELLER T I L T ON THE 1P AND 2P VIBRATORY STRAIN

60

1200 .0 7 i

1000.0 -

1000.0 -

800.0 - Jz \ E 0 [r u U E 600.0 I I I

800.0 r \ S

0 u H

a

600.0

MODEL PROP-FRN TESTS 1 .8 METER WIND TUNNEL

NRSR LEWIS w INBOARD BENDING S T R R I N B C 5 - 1

PlACH NO. -0.7 BLADE ANGLE - 59 DEG. PROPELLER SPEED - 6000 RPH 0 I/REV. STRAIN 8 2/REV. STRAIN

Y O O . 0 in

2 0 0 . 0

0 . 0

0.0 1.0 2.0 3.0 11.0 5.0 6.0 7 .0 PROPELLER T I L T ANGLE - DEG

PlFlCH NO. 1-7 ' BLADE RNGLE - 59 DEG. PROPELLER SPEED - 6500 RPH

1/REV. STRRIN 0 2/REV. STRAIN

c v,

900.0

200 .0

0 . 0

0.0 1.0 2 . 0 3.0 q.0 5.0 6.0 7 . 0 PROPELLER T I L T ANGLE - DEC

FIGURE 17. THE EFFECT OF PROPELLER TILT ON THE 1P AND 2P VIBRRTORY STRRIN

61

MODEL PROP-FAN TESTS 1.8 METER U I N D TUNNEL

NASA L E W I S w

1200.0 :

1000.0 1

800.0 1 S \ r 0 a

INBOARD BENDING S T R R I N B G 5 - I

800.0 - II \ r 0 cc V U

600.0 - I

MACH NO. -0.8 BLRDE ANGLE - 6 1 DEG. PROPELLER SPEED - 6000 RPM

1/REV. STRRIN o Z/REV. STRRIN

3 1000.0

L, 400.0

200.0

0.0

BLRDE ANGLE - 6 1 OEG. PROPELLER SPEED - 6500 RPM

0 l/REV. STRRIN o Z/REV. STRAIN

0.0 1.0 2.0 3 . 0 q.0 5.0 6.0 7.0 PROPELLER T I L T ANGLE - DEC

FIGURE 20. THE EFFECT OF PROPELLER TILT ON THE 1P AND 2P VIBRATORY STRAIN

62

MACH = 0.36

150 0 Case No. 1

r - - . - - ..

150

100 MICRO STRAIN PER EF

50

0

MACH = 0.6 r - - 6 Case no. 3

Case NO. 2 0 - *

.dQ - aG-L*g@35 " > :-- B,--ti:/-P- b

- - I I . I 4 4 I b m l I 1 1 1 " ' I . *

150

100 MICRO STRAIN PER EF

50

0

Figure 21 S t r a i n S e n s i t i v i t y v s . Power C o e f f i c i e n t f o r a l l Blade Angles, SR-3C-3 Prop-Fan t e s t s a t t h e 8 x 6 Wind Tunnel, NASA/Lewis. Inboard bending Gage 865-1. 63

c . MACH = 0.7 - o NASA LEWIS TEST 1

Case No. 40 I * 0 Q CALCULATED

e* G?

- Case No. 5 6 - - kc

4 arc' m-2 z 2 0 '0 y. - - -

I , . . I I f . l l , l I l l l I I 1

r

Lase Mu. sa@ L 0

STRA I N PER EF

MACH = 0.84

0

-0

STRAIN PER EF

50

STRAIN t U

-0 PER EF

50

* * e "" "" l " ' * " d

0.5 1.6 1-5 2.0 0- 0 CP - POWER COEFFICIENT

Figure 21 (Concluded)

S t r a i n Sens i t iv i ty vs. Power Coefficient f o r a l l Blade Angles, SR-3C-3 Prop-Fan t e s t s a t the 8 x 6 Wind Tunnel, NASA/Lewis. Inboard bending Gage BG5-1.

64

Micro Strain Per EF

Micro Strain Per EF

Micro Strain Per EF

150

100

50

0

150

100

50

0

150

100

50

0

MACH = 0.36

MACH = 0.70

Figure 22 Strain Sensitivity vs. Power Coefficient for all Blade Angles, several Prop-Fan response tests at the 8 X 6 Wind Tunnel, NASA/Lewis. Inboard Bending Gages.

65

Micro Strain

P e r EF

150

100

56

0

MGCH = 0.80

MACH = 6.84

Figure 22 S t r a i n S e n s i t i v i t y vs. Power C o e f f i c i e n t fo r a l l Blade (Concluded) Angles, severa l Prop-Fan response t e s t s a t the 8 X 6

Wind Tunnel, NASA/Lewis. Inboard Bending Gages.

APPENDIX I

P e a k T o t a l S t r a i n T a b u l a t i o n .

67/68

Y

L L 0 - W 0 e

c E

(0

* @I

2

t

a c P X

1.

c I- C c P w

r C C.1

(c 0 I- U LL: + w c L r_ 'A

i

-

N I

Q m ffi - I Q (1 m

m

(1 ffi

m

-4 I

m w n

- I m m L

N

m rt n

4 I

m P

r.i

N I

m P

r.1

m I

P P

3

r.i I

0 P

,+

,+ - I 0 LT1

* J 9

...... o m m ~ m m 0 1 a 0 1 o m c m m m o m m

......................... r( m O. n Y a Y a r.i- O N - Y Q o O. 'Q r\ in h a - h o n n o a Y O LO r.1 m Y O. O. m Y a o o O. h r d o o o Y) o t . i r . i ~ m m o o r ~ ~ moo om^‘^ - ~ ~ I I , Y Y -

.................................................. Q o NO. - o o o PO. ~li o m P (EA rl O. m m O. P r\ r.1 m 01 m ,+Y Y O o ' om Y o Q Y o m a Y o o n r.1 r.1 P - Q a N a - P r d o o VI Y - - o Q h h o I\ c - o n 10 Y o p\ h Q o P o. h h a o o O. n N m m Y o o o. Q N LI - - o o P n - - r.i o n m m m m N r+ r.i r d - N m m m m m rd r.1 N N m m m P o n - m m P o o n Q

69

.................................................. m o. n o rd o m o n P r.1 r( c. h rl m P m hi rl 01 a rl r( m -a N P o P N w m rd .-I m .om P h r.101 h m rl a .o P m m h P m P m m P r d o 01 UY m .o m o. m r( P w m m o. P m m m ~ i ) m P P P q m m Q w n n + 4 .o

m rl rl rl o. h o. P ~ o . ~3 r\ rl rl

LIY n I\ r d m w P P w o m li) in n w s h m n LY m m .o I\ .-I rl rl rl rl r(

70

O W ~ G ~ N A E PAGE rs Of POOR QUALITY

Q

LL 0

m

w e c iL

c S e +4

(0

* C4

-I-

rc c I X

.- < b-

d

2 G w

S 4 LI)

ic 0 c U b I-

: L '2 4

N I * 0 m

.A i * (1 m

m I

(1 m m

d I

m 0 P

.A I

m W P

N

m 0 m

d I N W I

c1 CJ W I

m I

e P

4

C l I

e P

d

3

* I

cl r=

* 0 9

......... ~ ) ~ . ~ h m t o . m o w o. N t m m N t r d l o t P m m.4 117 m m

..........

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APPENDIX I1

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