NASA/TM-2002-211894 

Methodology of Blade Unsteady Pressure 

Measurement in the NASA Transonic 

Flutter Cascade 

J. Lepicovsky 

QSS Group, Inc., Cleveland, Ohio 

E.R. McFarland 

Glenn Research Center, Cleveland, Ohio 

v.R. Capece 

University of Kentucky, Paducah, Kentucky 

T.A. Jett 

Glenn Research Center, Cleveland, Ohio 

R.G. Senyitko 

QSS Group, Inc., Cleveland, Ohio 

October 2002 

 ~_l 
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NASA/TM-2002-211894 

Methodology of Blade Unsteady Pressure 

Measurement in the NASA Transonic 

Flutter Cascade 

J. Lepicovsky 

QSS Group, Inc., Cleveland, Ohio 

E.R McFarland 

Glenn Research Center, Cleveland, Ohio 

V.R Capece 

University of Kentucky, Paducah, Kentucky 

T.A. Jett 

Glenn Research Center, Cleveland, Ohio 

RG. Senyitko 

QSS Group, Inc., Cleveland, Ohio 

National Aeronautics and 

Space Administration 

Glenn Research Center 

October 2002 
Acknowledgments 

The work was sponsored by the NASA Glenn Research Center under the Smart Engine Components Project 

managed by Mr. R.D. Corrigan as a part of the R&T Base Propulsion and Power Program. The authors would like 

to thank Mrs. L. Shaw, Chief, Turbomachinery and Propulsion Systems Division, for her continuous support. 

Trade names or manufacturers' names are used in this report for 

identification only. This usage does not constitute an official 

endorsement, either expressed or implied, by the National 

Aeronautics and Space Administration. 

The Aerospace Propulsion and Power Program at 

NASA Glenn Research Center sponsored this work. 

Available from 

NASA Center for Aerospace Information 

7121 Standard Drive 

National Technical Information Service 

5285 Port Royal Road 

Hanover, MD 21076 Springfield, VA 22100 

Available electronically at http: //gltrs.grc.nasa.gov 
.. __._~ ._ .. _ .._

METHODOLOGY OF BLADE UNSTEADY PRESSURE MEASUREMENT 

IN THE NASA TRANSONIC FLUTTER CASCADE 

Summary 

J. Lepicovsky 

ass Group, Inc. 

Cleveland, Ohio 44135 

E.R. McFarland 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

V.A. Capece 

University of Kentucky 

Paducah, Kentucky 42002 

T.A. Jett 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

R.G. Senyitko 

ass Group, Inc. 

Cleveland, Ohio 44135 

In this report the methodology adopted to measure unsteady pressures on blade surfaces in the 

NASA Transonic Flutter Cascade under conditions of simulated blade flutter is described. The facility was 

designed for this purpose, and previous test programs to study flutter have been conducted in the facility. 

Data from these previous programs proved to have some irregularities. To improve the quality of data 

taken in the facility, the cascade configuration was modified to improve the flow quality. Recent 

measurements show that the flowfield uniformity and blade loading periodicity were noticeably improved. 

Also, the facil ity was equipped with a completely new unsteady data acquisition system solely based on 

digital recording technology. New data acquisition and data reduction procedures were developed inhouse. The experimental data were reduced to pressure-time history segments for each blade and 

pressure port. The previous work done in this cascade, before its modification, reported that the 

oscillating cascade produced waves, which for some interblade phase angles reflected off the wind tunnel 

walls back into the cascade, interfered with the cascade unsteady aerodynamics, and contaminated the 

acquired data. To alleviate the problems with data contamination due to the back wall interference, a 

method of influence coefficients was selected for the future unsteady work in this cascade. In this 

approach only one blade in the cascade is oscillated at a time. 

The majority of the report is concerned with the experimental technique used and the 

experimental data generated in the facility. The report presents a list of all test conditions for the small 

amplitude of blade oscillations, and shows examples of some of the results achieved. The report is not a 

complete archival collection of unsteady pressure data acquired during this phase of experiments. All the 

acquired unsteady pressure data will be available in a suitable media form later. Also in the report is a 

description of the method of influence coefficients used to analyze unsteady flows, and a short discussion 

of a simple quasi-unsteady CFD simulation of the experiment that proved useful in planning the test 

program. The report does not discuss data analysis procedures like ensemble averaging, frequency 

analysis, and unsteady blade loading diagrams reconstructed using the influence coefficient method. 

Finally, the report presents the lessons learned from this phase of the experimental effort, and suggests 

the improvements and directions of the experimental work for tests to be carried out for large oscillation 

amplitudes. 

NASAffM-2002-2 l1 894 1 
Symbols 

C [mm] blade chord 

cp [1] pressure coefficient 

Cp [1] influence coefficient 

g [m.s-2

] gravitational acceleration 

h [mm] blade height 

iFL [dg] flow incidence 

iGM [dg] geometry incidence 

k [mV.kPa-'] slope of a calibration line 

Ma [1] flow Mach number 

p [kPa] unsteady pressure 

Pw [kPa] wall static pressure 

P1t [kPa] cascade inlet total pressure 

pp [kPa] pressure side pressure 

Ps [kPa] suction side pressure 

q [mV] y intercept of a calibration line 

s [mm] blade pitch 

V [m.s-'] flow velocity 

W [mm] cascade test section width 

x [mm] axial distance (cascade) 

y [mm] pitchwise distance (cascade) 

a [dg, rd] amplitude of blade oscillations 

y [dg] stagger angle 

7rTB [1] cascade pressure ratio 

p [kg.m-3] density 

e [dg] leading edge camber angle 

0' [dg] interblade phase angle 

Abbreviations 

AC alternating current 

ACM accelerometer 

AID analog/digital 

DC direct current 

EDM electro discharge machining 

OPP once per period 

NI National Instruments 

PC personal computer 

UDSA unsteady data acquisition system 

NASAfTM-2002-211 894 2 
1. Motivation 

Modern turbofan engines employ a highly-loaded, low-aspect ratio fan stage with transonic or 

low-supersonic relative velocities in the blade-tip region. The tip-section airfoils of these fan blades are 

noticeably different from the airfoil sections on the rest of the blade. The tip-section airfoils are designed 

for precompression, with a concave suction surface just downstream of the leading edge, and with very 

low overall camber. The airfoils have a sharp leading edge that makes them prone to flow separation at 

off-design conditions. Due to extreme flight envelope requirements, the engines are often operated near 

the stall flutter boundary of the fan. At these conditions the fan tip section is subjected to high incidence 

angles and high subsonic or transonic relative Mach numbers. Blade flutter and associated high cycle 

fatigue problems result. These are very detrimental to the engine health and must be avoided. Stall flutter 

and particularly blade life prediction codes are not yet fully reliable; their verification is hampered by a lack 

of reliable unsteady loading data, particularly for the airfoils in question. Interest in fan blade stall flutter 

research has increased in recent years. 

The NASA Transonic Flutter Cascade has been used to investigate behavior of a cascade of 

modern, low-aspect ratio fan blades operating near the stall flutter boundary that occurs at high incidence 

angles and high subsonic and transonic relative Mach numbers. A view of the cascade test section is in 

Fig. 1.1 . The blades for this experiment were designed and manufactured by Pratt & Whitney Engine 

Company. The airfoil and cascade parameters are given in Fig. 1.2 and Tab. 1.1 (Ref. 1.1). Previous 

measurements on these blades were reported by Buffum et al. (Refs. 1.1 and 1.2) at Mach numbers 

between 0.5 and 0.8, and chordal , geometric incidence angles, iGM , of 0 dg and 10 dg. Experimental 

blade-surface pressure distributions for steady flow at inlet Mach numbers of up to 0.5 were compared 

with various CFD predictions, but good agreement was found only up to 85% of chord and only for the 

incidence angle of 0 dg. The cascade flow periodicity was measured only for the three middle blades and 

was found sufficient for unsteady surface-pressure data sets acquired at full-chord reduced frequencies of 

0.4 and 0.8, and for an interblade phase angle of 180 dg. 

The unsteady data, reported by Buffum et al. (Refs. 1.1 and 1.2), were acquired for oscillations of 

all nine blades. In older work done in this cascade Buffum and Fleeter reported (Ref. 1.3) that the 

oscillating cascade produced waves which for some interblade phase angles reflected off the wind tunnel 

walls back into the cascade and interfered with the cascade unsteady aerodynamics. Later, the tunnel 

was provided with perforated walls at several locations and acoustic treatment. The efficiency of this 

arrangement still has not been proven. Ott at al. (Ref. 1.4) recommended for their facility just the 

opposite: replacement of slotted walls with solid ones. 

To alleviate the problems with data contamination due to the back wall interference, it was 

decided to adopt a method of influence coefficients for the future unsteady work in Transonic Flutter 

Cascade. The technique of influence coefficient technique is described in detail in Chapter 5. In this 

approach, only one blade in the cascade is oscillated at a time and the resulting unsteady pressures are 

measured on the remaining (nonmoving) blades. The unsteady aerodynamics of an equivalent cascade 

with all blades oscillating at a specified interblade phase angle is then determined through a vector 

summation of unsteady data from individual blades (Ref. 1.5). It must be stated that the use of the 

influence coefficient technique has only been demonstrated for attached flow. This facility has the unique 

capability of demonstrating this technique for separated flow. This technique requires a high degree of 

flow periodicity within the cascade for a larger number of blades. 

2. Cascade facility upgrade 

The NASA GRC Transonic Flutter Cascade is one of a very few test facilities dedicated to the 

unsteady aerodynamics of oscillating airfoils. The facility is used to provide data for modeling 

aerodynamics of blade stall flutter. The facility combines a linear cascade wind tunnel with a high-speed 

drive system that imparts pitching oscillations to cascade blades. The cascade consists of nine blades. All 

the blades or any single blade can be oscillated at realistic reduced frequencies (Strouhal numbers). 

NASAffM-2002-211894 3 
Interblade phase angles can be varied in increments of 15 dg. The facility has been described in detail in 

works of Boldman and Buggele (Ref. 2.1), Shaw et al. (Ref. 2.2), and Buffum et al. (Ref. 1.1). 

Improve~ cascade configuration 

Recently, a systematic experimental and computational study (Refs. 2.3 through 2.5) was carried 

out with a goal of improving the quality of experimental data acquired in the NASA Transonic Flutter 

Cascade facility. In the study, available means for control and modifications of the cascade flowfield were 

fully investigated and analyzed. Further, several discrepancies in the older data sets were explained, 

particularly the questions of actual flow incidence angles and the inconsistency between predicted and 

measured backpressure levels. Several conclusions important for future work on unsteady pressure 

distributions in a cascade under stall flutter conditions were reached. The most important conclusions can 

be summarized as follows: 

1. The boundary-layer bleed did not improve the flowfield uniformity and blade-load periodicity of the 

cascade tested. On the contrary, high bleed affected pitchwise distribution of flow incidence angles 

along the cascade and thus contributed to blade load variations. 

2. The boundary layer bleed improves flow uniformity in the spanwise direction only in the vicinity of 

both sidewalls. The midspan section of the blade is not affected at all. For future experiments, it is 

recommended that the boundary-layer bleed system not be used. 

3. The tailboard setting significantly affects the pitchwise distribution of static wall pressures and their 

levels. 

4. The overall turning of the flow by the tunnel walls must match the cascade blades ability to turn the 

flow when operating as an infinite cascade. A mismatch between the wall turning and the blade's 

turning of the flow will result in walls interfering with the blade-to-blade periodicity of the cascade. 

5. It is extremely important to match the tunnel wall contours with expected streamlines. This is 

particularly important for conditions of blade separated flows where it may not be readily obvious. Any 

mismatch between the cascade streamlines and tunnel contour walls results in a tunnel driven flow 

with the cascade under investigation having very little effect on the flowfield. 

6. Flow incidence angles should be measured for each new configuration of the cascade facility . It is 

very risky to rely on geometric angles in determining the actual flow incidence. 

7. By carefully tailoring tunnel wall contours to the expected streamlines for the current flow incidence of 

10 dg, the flowfield uniformity upstream and downstream of the cascade improved significantly. 

Further, very high blade load uniformity now extends over six blades from blade BL2 to blade BL7. 

A schematic picture of the modified cascade configuration is in Fig. 2.1. Wall static pressures 

distributions for this configuration were very uniform as shown in Fig. 2.2 for three inlet Mach numbers of 

0.5, 0.8, and 1. O. The saw tooth character of the upstream pressure distribution for Ma = 1.0 indicates 

presence of shock waves emanating from the blade leading edges. The uniformity of wall static pressures 

in the pitchwise direction was a good indicator of an improved cascade flow periodicity for this 

configuration . 

Instrumented blades 

Six blades were instrumented for this study. Two blades were instrumented with conventional 

static taps, 15 ports along the midspan line, one blade on the suction surface (blade S1) and the other on 

the pressure surface (blade P1). Figs. 2.3 and 2.4 show location of static ports on blades S1 and P1 . Four 

blades were instrumented with miniature high-frequency pressure transducers Kulite XCQ-062-15A with a 

nominal range of 0 to 100 kPa absolute. The blades instrumented with Kulite transducer were labeled 

KS1 and KS2 for instrumented suction surface, and KP1 and KP2 for instrumented pressure surface of 

the blade. The batch of transducers for blades KS1 and KP1 had an average sensitivity of 1.3 mVl kPa 

and the batch for blades KS2 and KP2 had an average sensitivity of 0.6 mVl kPa. Locations of Kulite 

transducers were identical to blades S1 and P1 respectively. The blades were grooved before the Kulite 

transducers were flush mounted in the blades. The machining was done by the EDM (Electric Discharge 

NASAfTM-2002-2 11 894 4 
Machining) method. The EDM electrode and the grooved blade are shown in Fig. 2.5. Fig. 2.6 shows the 

KS2 and KP2 blades prior inserting the Kulite transducers. Finally, Fig. 2.7 shows fully instrumented 

blades KS2 and KP1 . 

As seen in Fig 2.7, the wires from Kulite transducers were routed through the mounting shaft of 

the blade. This is a challenge because once the instrumented blade is inserted into the cascade, 

altogether 60 wires per blade (4 wires per each transducers) must be routed from the blade shaft pass 

the driving cams out of the cam gear box that is filled with oil. Fig. 2.8 shows the open cam box on the 

back side of the cascade test section and the details of transducer leads (see Ref. 2.1 for details about 

the blade drive system). The lead wires were terminated with miniature 32 pin connectors (two per each 

blade) that enabled connections with the data acquisition system. The Kulite leads are only 0.1 mm in 

diameter and quite fragi le. To prevent wire damage during blade installations and operation of the 

cascade, the wires were cemented into the blade grooves and also in the shaft. Unfortunately, it turned 

the weakest point was at the back face of the blade shaft, where the wires exited from the cement filler. 

Many transducers were lost due to wire breaking at the shaft backface. 

3. Data acquisition procedure 

Two sets of data were acquired during the experiments: steady state data and unsteady data. 

The steady state data were acquired using the labwide ESCORT system (Ref. 3.1). The ESCORT data 

system monitors and records steady state cascade research data as well as of the facility operation data. 

The unsteady data were acquired using a test cell dedicated system that was tailored to tne needs of this 

particular test setup. A requirement was to record the unsteady data in a digital form with a direct access 

to PC environment. 

Unsteady data acquisition system 

An overall diagram of the unsteady data acquisition system (UDAS) is shown in Fig. 3.1. The 

system was built around the National Instruments (NI) PXI Modular Instrumentation System. The PXI1010 system was selected because it offers rugged, shielded construction that provides a low-noise 

environment for data acquisition and signal conditioning. The core of the system is an embedded data 

acquisition computer PXI-8156B with a 333 MHz Intel Pentium processor and a 4 GB hard drive. Two NI 

6071 E AID (analog/digital) boards are controlled by the data acquisition computer. Each board accepts 

64 single-ended analog inputs, which means that up to 32 Kulite transducer can be connected to a board 

because each Kulite transducer requires a differential input (two channels per a transducer). The board 

has 12-bit resolution, it is 4096 discrete voltage levels. The scanning frequency is quite high 1.25 MHz, 

which allows sampling frequency for any of the 32 differential channels to go up to 39 kHz. The boards 

communicate with the data acquisition computer PXI-8156B via an internal bus at 1.25 MS/s. 

Each Kulite transducer had an allocated amplifier for signal conditioning. Endevco 4430A 

programmable bridge transducer signal conditioners were selected. This conditioner is a precision DC 

bridge amplifier with built-in constant excitation voltage/current supply and a programmable low pass 

filter. All the Kulite signal amplifiers were set for DC amplification of 50 for blades KS1 and KP1 , and 100 

for blades KS2 and KP2. Low pass filters and zero offsets were disabled. An oscilloscope Tektronix 

2465B was used to monitor live signals. 

The UDSA system was controlled by in-house developed software. All data acquisition and 

reduction software procedures were written in programming language G (NI Labview). Data acquisition 

time for each test point was 3150 ms, which is 120,000 samples. Raw data for each blade were stored in 

binary files in a spreadsheet form with 16 columns (lead signal plus 15 pressure transducers). The fi le 

size is 7.5 MB. In order to maintain the high speed of data sampling of 38 kHz, the AID boards were 

operated one at a time. It means that for each cascade operating condition to sets of data were acquired. 

NASAfTM-2002-211 894 5 
First one recorded board #1 data (blades KS1 and KS2), and the second one recorded data for board #2 

(blades KP1 and KP2). 

Data acquisition timing 

The timing of data acquisition is depicted in Fig. 3.2. It is identical for both AID boards. In this 

diagram, the elapsed time is shown on the horizontal axis, and the sequence of 32 channels per board on 

the vertical axis. Starting in the lower left corner, the first signal to acquire is the once per period (OPP) 

signal, which is once per blade oscillation period mark. The OPP signal is the lead signal for blade #1. 

The second signal to be recorded is the transducer #1 on blade #1 (KS1 for board #1 or KP1 for board 

#2). The time difference between these two signals is 0.8 which corresponds to the 1.25 MHz scanning 

frequency. The signals from the rest of the transducers on the blade are recorded in the same sequence. 

As seen in the picture, it takes 11 .2 ps to record data from all the transducers on one blade. Channel 17 

on this board records ACM signal, which is the accelerometer signal located on the tip shaft of the blade 

BL5, it is the blade that is oscillated. The purpose of this signal is to verify the position of the oscillating 

blade with respect to the recorded pressure time histories on the instrumented blades. The ACM signal is 

the lead signal for blade #2. Signals from pressure transducers on blade #2 are recorded in the same 

sequence as on blade #1. It takes all together 24.8 ps to record pressure data from both blades including 

OPP and ACM signals. The remaining 1.5 ps the system idles to complete the time period of 26.3 ps, 

which corresponds to the sampling frequency of 38 kHz. This cycle is repeated 120,000 times to 

complete one test point. There were two accelerometers available for some of the runs. For those test 

points, the accelerometer ACM_1 was allocated to board #1 and accelerometer ACM_2 to board #2. For 

the OPP signal however, it was the same signal that was always allocated to both boards. This signal 

serves as a synchronizer for the time history signals from two separate boards. This allocation scheme 

was maintained during the entire test program . Even if a transducer was lost, the timing sequence stayed 

the same, and the fau lty transducer data was removed only during the postprocessing procedures. 

Blade 'zero' angle 

Beginning of data acquisition is triggered randomly, and therefore the knowledge of the phase 

between the OPP mark and the blade angular position is important to establish zero for blade oscillations. 

Fig. 3.3 explains the phase relation between the oscillating blade and the OPP signal for the amplitude of 

blade oscillations of 0.6 dg. In the upper left corner, the blade is at 30.0 dg with respect to the horizontal 

plane. This is the blade 'zero' angle. During one period of blade oscillation, the blade angle first increases 

to 30.6 dg, then decreases to 29.4 dg, and returns back to 30.0 dg. The blade oscillation is forced by 

rotation of a camshaft (Fig. 2.8). The gear ratio is 6:1, it means that for one cam shaft revolution the blade 

oscillates six times. The cam shaft is provided with a hexagon contour section at the very end. A proximity 

capacity sensor faces the hexagon surface and generates OPP signal. Time history of the OPP signal for 

blade oscillating frequency of 200 Hz is shown at the bottom of Fig. 3.3. In this diagram, the interval 

between two adjacent peaks corresponds to one blade oscillation period, interval between every sixth 

peak corresponds to one camshaft revolution. The angular position of the proximity sensor was arbitrary 

with respect to the hexagon surface but stayed the same for the entire test program. Rotating the 

camshaft to the nearest hexagon vertex (to the nearest peak in time history plot) causes blade to move to 

an angle of 30 dg 31 min. From the above data one can determine the blade 'zero' angle instant in the 

OPP signal time history. For example, for 200 Hz oscillating frequency, the blade passed through the 

'zero' angle 1.08 ms past any peak instant in the OPP time history. Because the time history of the OPP 

signal is firmly connected with the time history of any signal on the same AID board by the board 

hardware clock (Fig. 3.2), the 'zero' blade position can be uniquely determined for any pressure history 

data. The direction of cam rotation during the pressure data acquisition is indicated in Fig 3.3 by a 

contoured arrow. 

NASAfTM-2002-211 894 6 
Transducer calibration 

A general impression is that the miniature pressure transducers, like Kulites, are suitable mainly 

for measurement of pressure fluctuations and the steady state pressure levels are not very reliable. 

Usually, AC amplifiers amplify signals from Kulite transducers, and the information about the 'DC' 

pressure level is then lost. This approach evolved in the past due to a low resolution of the AID boards 

used in data acquisition systems. The range and resolution of the NI AID board 6071 E used in the UDSA 

system, however, is large enough to record the absolute output voltage using DC amplifiers. The resulting 

pressure resolution is better than 24 Pa, which is fully sufficient. For comparison, the labwide steady state 

data recording system ESCORT has a resolution of 20 Pa for 100 kPa transducers. By recording absolute 

voltage output one can retrieve the information about steady-state pressure values as well as about the 

pressure fluctuations. By comparing the average absolute pressure levels with levels recorded by 

conventional static taps, it is possible to assess reliably the accuracy of measured pressure fluctuations. 

Extensive calibration tests of transducers embedded in the instrumented blades were carried out 

in a vacuum chamber. The tests were time consuming because the blades had to be dismounted from the 

cascade for each calibration test. Only one blade could be tested at a time. As an example of calibration 

results, Fig. 3.4 presents data blade KP1 , ports 6 and 10. The figure shows the calibration curves and 

deviation scatter. The results of four calibration tests of blade KP1 are summarized in Tab. 3.1 . As seen 

here, the calibration constants drift a little. The slope of the calibration curves is relatively stable; the 

scatter is about 0.2% of the fullscale (FS) value, which is 200 Pa. The scatter of zero offsets (y_O, y 

intercept) shows a larger value, up to 0.6% of FS (600 Pa) . To improve the accuracy of voltage/pressure 

conversion for Kulite transducers the transducers should be calibrated just prior each test run. This was 

impractical so for the slope value the results of the last calibration tests were used. For the zero drift, 

however, new 'zeros' were read prior each test and used for the pressure conversion. In this way, the 

accuracy of the steady state pressures (time average pressures of Kulite readings) improved noticeably. 

As an illustration, Fig. 3.5 shows a comparison of five tests of three blades. The plots are in a form of 

pressure coefficient (cp) distributions over a suction surface of the cascade blade BL4 for the inlet Mach 

numbers 0.5 and 0.8. The data collapse between the conventional tap data and four Kulite data is very 

good; the largest difference appears to be for port 6. In terms of the pressure accuracy the maximum 

difference between the static tap and averaged Kulite data for port 6 is less than 2% of the local dynamic 

pressure value for Mach number 0.5, and about 5% of the local dynamic pressure value for Mach number 

0.8. In terms of absolute local pressure values the tap/Kulite agreement is 0.4% for Mach number 0.5, 

and 3.4% for Mach number 0.8. The good agreement of Kulite averaged pressure values with the static 

tap values is a measure of accuracy for pressure fluctuations measured by Kulites. It is estimated that the 

accuracy of the pressure fluctuation amplitude measured by Kulite transducers (the AC component of the 

signal) is better than 2% of the local amplitude value. 

4. Data reduction procedures 

Signals from all pressure transducers were digitized and recorded as voltage levels. A low 

voltage unsteady signal is very sensitive to contamination due to electric ground loops and radio 

frequencies in general. All precautions were made to eliminate ground loops and properly shield all signal 

carrying wires, nevertheless some signals exhibited contamination and spurious spikes that could not be 

associated with flowfield pressure changes. Raw voltage signals were first converted to pressures, then 

in-house developed data cleaning procedures were applied to remove obvious signal contamination , and 

finally pressure levels were corrected for transducer zero drifts. All steps during the data reduction 

process were recorded in the data file header, and the original 'uncleaned' data are also available for 

comparisons. 

Data reduction process consisted from four procedures. They were: 

a. Voltage / pressure conversion, 

b. Data clipping, 

c. Data patching, 

d. Mean pressure correction. 

NASAfTM- 2002-2 11 894 7 
L 

Voltage/pressure conversion 

The voltage/pressure conversion procedure performs several functions . First, it retrieves voltage 

data for a selected port from a binary raw data blade file. Then, it converts voltage data to pressure units 

using the calibration constants as recorded in the latest calibration data file for a selected port. Next, it 

generates time variable based on the selected sampling frequency. Finally, it stores the pressure data as 

a function of time (pressure history) in a ASCII data file, that is labeled to indicate cascade configuration 

(blade position), blade label, port number, and test number that reflects operating conditions (data 

labeling convention is explained in Chapter 7). The output files are identified by an extension "hr3". The 

length of the entire record is 3000 ms. Fig. 4.1 shows a 1000-ms and a 100-ms intervals of a raw 

uncontaminated signal. 

Data clipping (signal local overshoots, single point spikes) 

The signal clipping procedure is based on the following observation. Visualization of the cascade 

flowfield revealed noticeable flow instabilities for inlet Mach numbers in a range from 0.93 to 1.03. For 

these inlet flow conditions, the shock wave pattern in the cascade is very unstable and exhibits rapid 

switching among individual blade channels (Ref. 4.1 ). This flow instability is strongest on the blade 

suction side close to the leading edge. Switching between the supersonic and subsonic flows in this 

region causes rapid changes in the local static pressure in the same region. The static pressure is at a 

low level for supersonic inlet flow and jumps to a high-pressure level when the flow drops to a subsonic 

velocity. Fig 4.2 shows the pressure jumps as recorded at port 1 for an average inlet flow at Mach number 

of 0.9 7. As seen here, the rapid pressure changes are of the order of 20 to 25 kPa. Even that these 

pressure jumps are the fastest pressure changes in the cascade, they do not happen instantly but over an 

interval of 300 to 350 J.1s. As discussed above, the pressure signal was sampled at 38 kHz, which means 

that two adjacent data points are 26.3 J.1s apart. In other words , the fastest pressure changes are 

stretched at least over 12 sampling periods. Therefore, a pressure change between any two consecutive 

samples cannot be larger than 1.6 kPa, maybe 2.0 kPa at maximum. Data clipping procedure was built on 

this fact. The procedure marches through the recorded data and examines every sequence of three 

consecutive data points. If the middle point exhibits pressure overshoot above or below a selected value, 

the data point is removed and replaced by an average of it's pre- and post-neighbors. It should be 

stressed here, that the clipping procedure does not affect the timing of the data sequence and corrects 

only a single point over or under shoots. An example is shown in Fig. 4.3a,b. The data files that were 

modified by clipping have an extension "hc3". 

Data patching (temporarily signal dropouts) 

"Single point" over- or undershoots, however, were not the only data contamination observed. 

Some of the transducers exhibited occasionally behavior resembling a temporarily open circuit 

overcompensated by an amplifier feedback resulting in saturation of the AID board input. An example of 

such behavior is in Fig. 4.4. These signal 'drop-outs" sometimes lasted over several sampling periods. 

Reasons for this behavior are still unclear, perhaps momentary disruptions of continuity either on the 

transducer bridge or in lead wires and connectors. Because the signal was contaminated over several 

sampling periods, the clipping procedure does not clean the signal, as demonstrated in Fig. 4.5. The 

signal outside of the saturation instants appeared healthy, and therefore a procedure was established to 

recover the undisturbed portion of the signal. The patching procedure is based on an assumption that a 

signal that represents true pressure cannot exceed minimum and maximum thresholds that defined a 

band about a mean pressure level. Fig. 4.6 illustrates this assumption. If the signal exceeds the selected 

thresholds, then the portion of the signal above the upper threshold (or below the lower threshold) is 

removed and the missing portion is replaced with a straight line connecting the endpoints of the patch, as 

shown in Fig. 4.7. This procedure, similarly as data clipping, does not alter the timing of the data 

sequence. This procedure, however, will slightly affect the variance of the signal. The saturation drop-outs 

extend only over few percent of the data collection interval. It is believed that the effects on the variance 

NASAffM-2002-211 894 8 
of the undisturbed signal are insignificant. The data files that were subjected to patch procedure have 

extension "hp3". 

Zero drift correction 

The last procedure corrects the pressure values for transducer zero drift that may be different 

from the calibration data. Before each test, zeros for all Kulite transducers were recorded. These 

readings, instantaneous zero offsets, respectively their differences from the zero offset recorded during 

the last previous Kulite calibration, were applied to the pressure values stored in the last modified data file 

(either "hc3" or "hp3" files). This procedure affects only the pressure average value (DC component of the 

signal). Magnitude of pressure fluctuations (the AC component) is not affected at all. The output file from 

the mean pressure correction procedure was labeled "hx3". These are the files that should be used for 

additional data analysis including ensemble averaging, frequency analysis, unsteady blade loading, 

influence coefficient method, etc. 

5. Influence coefficient method 

In the previous cascade experiments by Buffum, Capece, King, and EL-Aini (Refs. 1.1 and 1.2) 

only an interblade phase angle of 180 dg was reported. Experimental results at other interblade phase 

angles were judged to be unrealistic due to tunnel sidewall reflections. To expand the range of interblade 

phase angles possible, the influence coefficient method will be explored. 

The influence coefficient method is based on the superposition principle for linear systems. For 

this case the unsteady pressures on an airfoil for a given cascade geometry, reduced frequency, and inlet 

flow condition are separated into a contribution of the airfoil oscillating on itself and a separate 

contribution from the oscillation of each neighboring airfoil. Hence, the unsteady aerodynamic 

perturbations generated in a cascade with all airfoils oscillating can be represented as the sum of 

perturbations generated by oscillating the airfoils individually. The situation where all airfoils are oscillating 

at a constant interblade phase angle is typically referred to as the traveling wave method. 

The influence coefficient method has been used numerically and experimentally. For example, in 

a three dimensional numerical example for a compressor blade in supersonic inviscid flow, Gerolymos 

(Ref. 5.1) has shown that the influence coefficient method using 7 blade passages gives nearly equivalent 

first harmonic pressure coefficient results as the traveling wave method. 

Many investigators have conducted experimental studies directed at validating the influence 

coefficient method. Some detailed studies have been conducted by Buffum and Fleeter (Ref. 5.2), Bolcs, 

Fransson, and Schafli (Ref. 5.3) and Fransson (ref. 5.4). Buffum and Fleeter (Ref. 5.2) used biconvex 

airfoils. Bolcs et al. (Ref. 5.3) and Fransson (Ref. 5.4) used an annular cascade with a variety of different 

turbine configurations. In general, the results have indicated the validity of the influence coefficient 

method for attached flow for cases where the wind tunnel sidewalls for linear cascade have a negligible 

influence and for unsteady flows that are not in the vicinity of acoustic resonance. 

Theory 

The influence coefficient method can be used for any of the oscillating airfoil response 

coefficients, e.g. unsteady aerodynamic lift, unsteady aerodynamic moment, or unsteady surface 

pressure. For the present investigation the first harmonic unsteady surface pressure coefficient is of 

fundamental interest. 

NASAffM-2002-21 1894 9 
I 

I 

Mathematically, the relation between the unsteady surface pressure coefficient influence coefficient 

and the unsteady surface pressure coefficient for all airfoils oscillating at a constant interblade phase angle 

(cr) is given by Eq. 1. 

Where 

C/(x) 

N 

N 

C p (x, eJ ) = L C / (x)eil/O' (1) 

I/=-N 

is the complex value unsteady surface pressure coefficient at a given chord location (x) acting 

on the reference airfoil with all airfoils oscillating at a constant interblade phase angle (cr); 

is the complex valued unsteady surface pressure influence coefficient at the same chord 

location, which is the influence that oscillations of airfoil n have on the unsteady surface 

pressure coefficient of the reference airfoil; 

is the number of airfoils summed. 

For an infinite cascade N is infinity. In the present case of the NASA Transonic Flutter Cascade, 

which has nine airfoils, N will be a maximum of four. Note that this relation has the form of a Fourier 

transform. 

The influence coefficients can be determined from mathematical models for an infinite cascade of 

airfoils oscillating with a constant interblade phase angle. For this case the mathematical relationship is 

determined by inverting the Eq. 1. 

C/ (x)= _l_ r Cp(x, eJ )e-iIlO' deJ 

27r Jr 

(2) 

This relation can be used to determine analytical influence coefficients. Substitution of the 

analytical influence coefficients into Eq. 1 yields analytical results for a finite number of airfoils oscillating 

in an infinite cascade. 

Experimental procedure 

The experimental procedure to determine the unsteady surface pressure influence coefficients 

will now be discussed. The numbering system is illustrated in Fig. 5.1 for an infinite cascade where the 

"center" airfoil has been selected as the reference airioil, n = O. The influence coefficients are determined 

in a sequential manner. To determine the influence of the reference airfoil oscillating on itself Cp"=o (x) , 

the reference airfoil is oscillated with all other airfoils held stationary, as depicted schematically in Fig. 5.2. 

In an analogous manner, the influence of oscillating airfoil n = + 1 with all other airfoils held stationary on 

the reference airfoil is determined, C p"=1 (x). This procedure is repeated for as many other influence 

coefficients as desired. The unsteady surface pressure influence coefficients are then substituted into 

Eq. 1 to determine the unsteady suriace pressure coefficients at a constant interblade phase angle in the 

traveling wave mode. 

To validate the influence coefficient method experimentally the cascade will be operated at 

interblade phase angles where the effect of acoustic wall reflections on the measured unsteady pressures 

NASAfTM-2002-211894 10 
is negligible. This will be quantified by verifying the dynamic periodicity of the airfoils in the central portion 

of the cascade. The influence coefficient technique will be used with Eq. 1 and the results correlated with 

the experiments performed with all airfoils oscillating at a constant interblade phase angle. An oscillation 

amplitude of 0.6 dg will be used. For this small of an oscillation amplitude the unsteady flow is assumed 

to be linear. This assumption will be evaluated in this investigation. To quantify the influence of oscillation 

amplitude these experiments will be repeated for an oscillation amplitude of 1.2 dg, and selected 

experiments will be done at an oscillation amplitude of 2.4 dg. 

By using the influence coefficient technique interblade phase angles will be achievable that were 

not previously possible. This will allow the attainment of experimental data to further our understanding of 

the flow physics for unsteady separated flow. 

To complete this section an alternate method for obtaining the influence coefficients using two 

instrumented airfoils is illustrated schematically in Fig. 5.3. This arrangement will yield the n = + 1 and 

n = +2 influence coefficients. It can be seen that other influence coefficients can be easily obtained by 

moving the instrumented airfoils to other positions in the cascade. An attractive feature of this procedure 

is that it keeps the oscillating airfoil in the center of a finite cascade, which will minimize any acoustic 

wave reflections from the cascade sidewalls. In an infinite cascade this method is satisfactory and all the 

influence coefficients can be obtained by oscillating a single airfoil. However, for a finite number of airfoils 

there will be concerns with airfoils that are in the vicinity of the walls. In this region the airfoils will not be 

behaving as if they are in a cascade. Hence, the oscillating airfoil and instrumented airfoils will be shifted 

within the cascade to assess this difficulty. 

Analytical influence coefficients 

To complete this chapter analytical influence coefficients will be determined. The analytical 

influence coefficients will indicate how fast the influence coefficients decay as the oscillating airfoil moves 

away from the reference airfoil. 

The mathematical model used for the infinite cascade is the flat plate semi-analytical analysis 

given by Smith (Ref. 5.5). This model is for two dimensional inviscid compressible flow. For this example 

the present cascade configuration is used. Since the model is inviscid the flow will be attached, and the 

incidence angle is zero; i.e., no steady aerodynamic loading. A Mach number of 0.5 is used with a 

reduced frequency based on semi-chord of 1. The cascade stagger angle is 60 dg. 

The magnitude of the unsteady surface pressure difference influence coefficients determined 

using Eq. 2 are presented in Fig. 5.4. The influence coefficients are determined for N = 5. Note that the 

largest influence coefficient is for n = 0, the influence of the reference airfoil oscillating on itself, 

!1C /=0 ex / c) . The next largest influence coefficient is for n = -1, with a rapid decay in the influence 

coefficients as the airfoil that oscillates moves further from the reference airfoil. For this example, the 

influence coefficients are mostly negligible for Inl > 2. 

These influence coefficients are now summed using Eq. 1 and compared to the constant 

interblade phase angle calculation for interblade phase angles of 0 dg and 180 dg. For an interblade 

phase angle of 0 dg acoustic wave are propagating away from the cascade into the far field. The 

summation in Eq. 1 was taken to be for N = 2, and N = 5, and the results are presented in Figs. 5.5 and 

5.6. These results indicate that in both cases the influence coefficient series is rapidly convergent and 

that the reference airfoil and its immediate two neighbors have the most significant effect on the unsteady 

pressure difference coefficient. Furthermore, the influence coefficient summation exhibits excellent 

correlation with the traveling wave method. Moreover, the summation with N = 2 was sufficient except 

near the leading edge (where there is a square root singularity in the pressure distribution) for (Y= 0 dg. 

NASAffM-2002-211 894 11 
6. Simplified CFD analysis 

When improving the flow quality of the facil ity, a two dimensional computer simulation of the test 

section was developed. The integral equation method described by McFarland (Refs. 6.1 and 6.2) was 

used. The simulation included the nine cascade blades and the tunnel walls. Since the solution technique 

does not require a computational grid and is very fast running, it provided a means for quickly looking at 

the effects variations in the test section geometry had on the periodicity of the flow. Figs. 6.1 a and 6.1 b 

illustrate the simulation of the test section. When combined with more sophisticated Reynolds averaged 

Navier Stokes CFD methods of Chima (Refs. 6.3 and 6.4), the CFD simulation proved very useful in 

achieving the goal of improving the flow quality and in understanding the test section flow. 

Early in the unsteady test program, confidence in the unsteady instrumentation had not been 

established and the response of the flow field to the blade oscillation was unknown. When no response 

was initially being measured on the leading edges of the blades located three blades away from the 

central oscillating blade, concerns were raised that the instrumentation was faulty or that the experimental 

technique of oscillating a single blade was not good. 

In order to resolve these concerns , a simple modification of the integral equation method was 

developed that produced a quasi-unsteady simulation of the oscillating flow. This was accomplished by 

adding the equation of motion of the oscillating blade. This allowed the position of the blade and its 

velocity relative to the stationary blades and tunnel walls to be calculated for any arbitrary point in the 

oscillation cycle. This information was used to alter the previous stationary blade simulation by modifying 

the surface velocity boundary conditions according the blades position during an oscillation. The result 

was an instantaneous look at the potential effect the blade oscillation had on the flow field. 

The results of the calculations gave the location and the magnitude of variations in the flow field 

caused by the moving blade at a specific position in the oscillation cycle. A typical sequence of solutions 

is shown in Figs. 6.2a, band c. The test condition for the figures is inlet Mach number of O.B, oscillating 

frequency of 500 Hz, and maximum deflection of 0.6 dg. The first figure shows the blade at maximum 

deflection, leading edge moved 0.6 dg, to the right and at zero rotational velocity, as the blade is about to 

change direction. The middle figure shows the flow field with the blade at 0 dg deflection, but at maximum 

rotational velocity. The last figure shows the blade at maximum deflection to the left. Timing or phasing 

information cannot be determined from the solutions since time is not a variable in the solution. Any 

potential effect appears to occur through out the flow field instantaneously in the calculation. 

The entire test matrix for the two subsonic inlet Mach numbers of 0.5 and O.B were calculated. 

Four frequency (200, 300, 400 and 500 Hz) and three deflection angles (0.6, 1.2, and 2.4 dg) were 

calculated for these Mach numbers. A summary of the results of these calculations are presented in 

Fig. 6.3. 

Finally animations were made using the simulations. A series of solutions were made for the 

blade as it moved through one oscillation. Contour plots similar to those shown in Fig. 6.2 were made for 

each blade position in the series. When these are shown in rapid sequence the contour lines appear to be 

in motion. The locations in the simulation where the lines appear to be moving the most highlight the 

regions where unsteady flow variations will occur. These animations proved very useful in determine 

where to expect a high amount of signal response from the Kulite sensors, and also regions where the 

response from the sensors will be little. The animations also indicated test conditions where reflex of the 

unsteady pressure waves by the wall may occur. 

NASAffM-2002-211 894 12 
7. Samples of experimental data 

Test matrix 

Two sets of unsteady pressure data were acquired during the first phase of this experimental 

work. In the first set no blade in the cascade was oscillated and the inlet Mach number was gradually 

raised from 0.4 to 1. 1. In the second set, the middle blade (blade BL5) was oscillated at frequencies 200, 

300,400, and 500 Hz for the inlet Mach numbers of 0.5, 0.8, 1.0, and 1. 1. Blade BL5 was oscillated at a 

single amplitude of 0.6 dg. A simplified test matrix is shown in Tabs. 7.1 and 7.2. All test conditions for 

nonoscillating blade data are listed. For the oscillating blade data, the actual test matrix consists of all 

possible combinations of test parameters listed in Tab. 7.2. 

For both data sets, the blade BL5 was always the uninstrumented one. The Kulite instrumented 

blades KS1 , KS2 (suction side). and KP1 , KP2 (pressure side). as well as the blades instrumented with 

conventional static taps S1 (suction side), and P1 (pressure side) were randomly placed in blade 

positions BL2, BL3, BL4, BL6, BL7, and BL8 for recording pressure data throughout the cascade. 

Tab. 7.3 presents a list of blade positions for this phase of the experiment. The middle blade (BL5) is 

labeled FB (flutter blade). The last column designates the data group. This is the key to allocate the 

instrumented blade positions to individual test runs. 

Data labeling convention 

The following labeling convention was adopted for data assignment: 

Here, the first letter (C) denotes cascade data; the following number denotes cascade configuration (e.g. 

n = 1 means incidence 10 dg and oscillation amplitude 0.6 dg). Next, the data group is listed (e.g. gg = 35 

designates test group run on 11-Apr-OO when the blade KS1 was in blade position BL 4 as can be found 

in Tab. 7.3). This is followed by a test run sequential number (e.g. tt = 14). The test sequential number 

has no special meaning except of distinguishing individual test runs. The following group (Kbb) 

designates Kulite instrumented blade (e.g. Kbb = KS1 , means blade #1 instrumented with Kulite 

transducers on the suction surface). Each instrumented blade had 15 pressure ports as shown in 

Figs. 2.3 and 2.4. The pressure port is given next (e.g. pp = 03, means third pressure port counted from 

the blade leading edge). The file name extension has the following meaning: the first letter (h) denotes 

time-history data, the next letter (~ determines the type of data as far as data reduction and cleaning 

procedures are concerned (e.g. r = raw data, c = clipped data, p = patched data, and x = zero correction 

data). Finally, the last digit (/) determines the time interval of the pressure unsteady data (e.g. 3 = 

3000 ms of continuous data). In general all data that will be released to the public domain for additional 

data reduction, including conditional sampling and frequency analysis, will have extension hx3. 

Unsteady pressure data for no blade oscillations 

A serious of test data for blade KS1 and pressure port 01 for the case of no blade oscillation is 

shown in Fig. 7.1 . The figure shows a series of data records for an increasing inlet Mach number ranging 

from 0.5 to 1.02. Only the interval of 200 ms is shown for clarity. The presented pressure data are for 

pressure port 01 on blade KS1 (KS1_01 data files). Static pressures for the same port acquired by a 

conventional surface tap (blade S1) showed a continuous and smooth pressure drop through the 

transonic region as the inlet Mach number was raised. Such data indicate that the local flow velocity, just 

downstream of the leading edge, continuously increases from subsonic to low supersonic values. This 

indicated continuous velocity increase, however, contradicts the observed unsteady and intermittent 

behavior of the flow shock pattern for transonic inlet flow conditions (Ref. 4.1). Furthermore, surface flow 

NASAfTM- 2002-2 11894 13 

__ J 
--- . . _ 

visualization using a oil-paint mixture clearly shows that for the subsonic inlet Mach numbers there is a 

large separated flow region on the blade suction surface just past the leading edge exhibiting a complex 

three-dimensional flow structure (Ref. 4.1). For the supersonic inlet flow, however, the flow past the 

leading edge is fully attached to the blade on a considerable length. This abrupt change of surface flow 

patterns is also not indicative of a smooth velocity increase through the transonic flow region. 

Unsteady pressure levels at port 01 , presented in Fig. 7.1, show pressure behavior that is not 

indicative of a smooth velocity flowfield in the high subsonic and transonic flow regions, and perhaps give 

new insight in the physics of the subsonic to supersonic flow transition. For lower inlet Mach numbers (Ma 

< 0_85) the local flow past the blade leading edge is always subsonic with increasing velocity fluctuations 

in the separated flow region. Starting at this inlet flow Mach number, local pressure begins to drop 

temporarily below the level at which the flow reaches low supersonic velocities. These bursts of 

supersonic velocity are at first very short (a few milliseconds) and infrequent. With increasing inlet Mach 

number, however, the duration and number of appearances of supersonic flow velocities increases 

dramatically. For an inlet Mach number of about 0.95 to 0.97, the local flow velocity at the blade leading 

edge is supersonic half of the time. However, it appears that the velocity is still switching randomly and is 

of variable duration. As the Mach number increases further the regions of supersonic flow velocity rapidly 

lengthen with very sporadic instances of subsonic velocity pockets. 

The jumps in the local pressure level in the blade leading edge region are about 20 to 25 kPa for 

the Mach number range from 0.9 to 1.0. It appears that the flow in the high subsonic and transonic region 

is bistable, randomly switching between the subsonic and supersonic flows. The smooth pressure drop in 

this region measured by the conventional static taps is an artifact of the averaging process of this 

pressure measuring method. It appears that in reality the change in overall flow conditions in the 

transonic flow region is based on the frequency of switching between two stable flow states rather than on 

the continuous increase of the flow velocity. 

To the authors' knowledge this phenomenon of pressure bistability in the leading region of a 

transonic cascade was not reported yet. Pressure jumping between two levels in the transonic region 

generates large intermittent loading on the blade leading edge region and can lead to the onset of blade 

torsional oscillations resulting in severe flutter. At the moment, it is not clear if this is a real phenomenon 

in an annular transonic fan cascade or a 'byproduct' of a span restricted linear transonic cascade. The 

acquired data will be thoroughly analyzed before a final judgment can be reached. In any case, this new 

model of the flow physics in the transonic fan blade cascade will either impact computational methods for 

flutter predictions in transonic fans or limit the use of linear transonic cascade data to regions outside of 

the range of pressure bistable behavior. 

Unsteady pressure data for middle blade oscillations 

In this group of experiments the middle blade (blade BLS) was oscillated at frequencies 200, 300, 

400, and 500 Hz for the inlet Mach numbers of 0.5, 0.8, 1.0, and 1.1. Blade BLS was oscillated at a single 

amplitude of 0.6 dg. The following examples of unsteady pressure data are for blade KS1 located in 

position BL4, it is for n = -1 in terms of the influence coefficient method (see Chapter 5). 

The effects of blade oscillation frequency (blade BL5) are shown in Fig. 7.2. The segments 

100 ms long of pressure data for port 15 (the closest port to the airfoil trailing edge) are shown here for 

the case of no oscillations and oscillations of 200, 300, and 400 Hz for the inlet Mach number of 1.08. It 

should be stressed that the shown data are pure time histories, not enhanced by ensemble averaging. 

Several observations can be made here. First, the pressure signal for case of no oscillations 

(C1 _42_95_KS1 _ 1S) exhibits weak contamination due to modulation by the net frequency of 60 Hz. 

Peak-to-peak variation of the signal contamination is about 2 kPa. The distributions acquired for blade 

BL5 oscillating at 200, 300, and 400 Hz show very strong modulation due to the oscillating blade 

frequency. The peak-to-peak variations of the modulated signal are 9 kPa for frequencies of 200 and 

300 Hz (data files C1_ 42_96_KS1_15 and C1 _ 42_97 _KS1 _15), and about 7 kPa for the case of 400 Hz 

(data file C1_42_98_KS1_15). 

NASAffM-2002-211 894 14 

-----.. - ------

---I 
The situation for the subsonic inlet Mach number of 0.8 is depicted in Fig. 7.3. Here the signal for 

no blade oscillations (C1_ 42_32_KS1_15) exhibits very strong modulation at frequency of 91 Hz and 

peak-to-peak variation of 11 kPa. Pressure distribution acquired at port 01 on the same blade and the 

same inlet Mach number (data file C1_36_17 _KS1 _01 in Fig. 7.1 a) does not exhibit any periodic 

modulation at all. It is very doubtful that the observed pressure fluctuations at pressure fluctuations 

experienced by the trailing region of the airfoil for this flow conditions. It should be stressed that this is a 

preliminary conclusion and the data will be subjected detail analysis before a final statement can be 

made. Introducing blade oscillations of 200 Hz does not visibly change the character of the signal 

(C1 _42_33_KS1 _1S). For higher frequencies of blade BLS oscillations, the effects of blade oscillations 

are traceable (data files C1 _42_35_KS1 _15 and C1 _42_36_KS1 _15). In particular, for the highest blade 

frequency, a modulation of 500 Hz is superimposed on the signal and is traceable. The peak-to-peak 

variation of SOO-Hz modulation is about 5 kPa. 

Finally, data for the lowest inlet Mach number of 0.5 are shown in Fig. 7.4. The case of no blade 

oscillations shows contamination by the net modulation of 60 Hz (C1_ 42_ 13_KS1 _1S). The peak-to-peak 

variation is about 2 kPa. The effects of forced oscillations are barely traceable for the blade frequency of 

200 Hz (C1 _42_14_KS1 _1S), however, they are clearly visible for the cases of blade frequency of 400 

and 500 Hz (data files C1 _ 42_16_KS1_15 and C1_ 42_17 _KS1 _15). However, the peak-to-peak variation 

of the forced signal modulation is quite small, about 2 kPa. 

Changes in the character of the unsteady pressure signal from port to port along the blade chord 

are shown in Fig. 7.S for the case of a supersonic inlet Mach number of 1.08 and the forcing frequency of 

400 Hz. The first pressure distribution is for Port 01 , which is very close to the blade leading edge (data 

file C1 _42_98_KS1 _01). The signal is very clean, uncontaminated and correctly shows no effects of the 

oscillating blade BL5; for the supersonic inlet condition any pressure fluctuations generated at the leading 

edge of blade BL5 cannot propagate to the leading edge region of blade BL4. Signals from ports OS and 

06, however, already show effects of forced pressure modulation (data files C1 _ 42_98_KS1 _06 and 

CC 42_98_KS1 _07). The peak-to-peak modulations are about 3 kPa and 4 kPa, respectively. The Signal 

for Port 15 (data file C1_ 42_98_KS1 _15) was already discussed above, and shows the strongest 

modulation of 7 kPa peak-to-peak. 

It can be stated, based on this very short overview, that the presented pressure distributions for 

blade n = -1 appear to follow some of the predictions made using the simplified CFD analysis. The 

predicted results are summarized in Fig. 6.3. The discussed unsteady pressure distributions show that 

peak-to-peak variations of the forced modulation increase with an increasing inlet Mach number. For the 

blade suction side, the amplitude of forced modulation is larger for the pressure port at the blade leading 

edge than for the blade midchord region. Finally, there seems to be little effect of the forcing frequency 

level on the amplitude of forced modulation. Of course, these are only preliminary conclusions based on a 

visual inspection of the unsteady pressure Signals. A rigorous analysis, based on ensemble averaging 

and a frequency content of the experimental pressure data will reveal more details and reach reliable 

conclusions about forcing effects on unsteady pressures in a cascade of transonic fan airfoils. 

Once-per-period and accelerometer data 

The oscillating blade experiences variable acceleration. In an ideal case, the acceleration will 

follow a periodic sine function with zero values for the blade passing through the 'zero' angle position (it is 

at 30.0 dg angle). Consequently, measuring the acceleration function, synchronized with the pressure 

data acquisition, it should be possible to allocate easily the instant of the blade 'zero' angle in the 

pressure data history. An accelerometer cannot be placed on the blade while acquiring pressures 

because even the miniature accelerometers are 'bulky' and the blade and cascade flow patterns would be 

severely distorted. An attempt was made to measure the blade acceleration by placing two 

accelerometers out of the cascade flow, an arm that was attached to the blade-tip shaft that protruded 

through the cascade wall. The arrangement is shown in Fig. 7.6. Two Entran accelerometers model EGA5000 were used, but eventually only one accelerometer eventually worked. The span of the arm modeled 

the blade chord. The accelerometer was placed 40.2 mm of the axis of the blade shaft. The calculated 

NASAffM- 2002-211894 15 
maximum amplitude (at ±0.6 dg angle), velocity (at 0.0 dg angle) and acceleration (at ±0.6 dg angle) for 

the transducer location are given in Tab. 7.1 for the range of oscillation frequencies from 200 Hz to 

500 Hz. The accelerometer data recorded are shown in F~ . 7.7 together with the OPP signal. The figures 

show raw signals converted to 'g-values' (g = 9.81 m/s ) using the manufacturer stated sensitivity of 

0.061 mV/g. 

A first glance at the accelerometer data reveals that there are severe problems. It must be 

stressed here that the intention behind this exercise was to back up the OPP signal in determining the 

blade 'zero' angle using accelerometer signal, and not to measure accurately the blade acceleration. The 

accelerometers were used without prior calibration just to indicate the shape of the acceleration signal. 

The first problem is the level of the maximum accelerations that seems to be on the order of 60g to 70g 

for oscillation frequencies up to 400 Hz, which is wrong. In particular for frequencies of 200 Hz and 

300 Hz, this is not even possible because the maximum level should be 17g and 38g, respectively (see 

Tab. 7.4). The reason for that might be in the range of the available transducer, which was ±5000g. Even 

for a very high accelerometer accuracy of 1.0 "Ie , the maximum acceleration levels for 200 Hz and 300 Hz 

are below the accuracy threshold. For frequencies 400 Hz and 500 Hz, the recorded maximum 

acceleration reaches occasionally levels close to the expected values. The second problem is the shape 

of the acceleration function. Surprisingly, for the 200 Hz case, for which the maximum acceleration level is 

not reliable, the acceleration trace shape is quite clean . As indicated by the arrows in Fig. 7. 7a, the zero 

acceleration instants happen about 0.75 ms past the once-per-period peaks, which is close to the value of 

1.08 ms as determined from the static conditions described in Fig. 3.3. For the frequencies above 200 Hz 

the acceleration function is heavily contaminated with higher frequencies that do not allow determining 

instants of zero acceleration levels with an acceptable accuracy. The reason for this high-frequency 

contamination might be slipping of the arm on the shaft during oscillation , and insufficient arm stiffness 

that allow arm own oscillations (flexing). Both these problems will be addressed in the second phase of 

the flutter experiments that will be carried out for larger oscillation amplitudes. 

8. Lessons learned 

The following lessons were learned from the first phase of experiments on transonic flutter: 

Transducer leads. Four blades instrumented with 60 miniature Kulite pressure transducers were 

available at the onset of the research work. Unfortunately, half of the transducers were lost within first 

several test runs. The primary reason for it was cracking of transducer wires at the backface of the blade 

shaft. The thin transducer wires pass through the blade shaft and are spliced to thicker wires. The thin 

wires are cemented in the blade shaft. The thicker wires are routed through the cam box to get the 

transducer signals out of the cascade. The arrangement was illustrated in Figs. 2.7 and 2.8. Each 

transducer has four wires. If a single wire cracks , the transducer is lost. All damage to the wires 

happened just at the backface of the blade shaft. It is not feasible to repair individual leads cemented in 

the blade shaft. A better solution would be to enlarge the cavity in the blade shaft and spliced the fragile 

Kulite leads to thicker wires ahead of the shaft end , pull the splicing points inside the shaft and cement 

the thicker wires in the shaft end. This arrangement will give chance to repair broken wires by removing 

the cement plug, pull the thin wire out of the shaft, and redo splicing. Unfortunately, this is a very costly 

solution because the entire blade (15 transducers) must be reinstrumented with a new batch of 

transducers. 

Signal contamination. For operation reasons, the signal conditioners had to be located in the 

control room and not in the test cell. Low voltages from Kulite transducers were carried over a distance of 

12 m. All precautions were made to eliminate ground loops and properly shield all signal carrying wires, 

nevertheless some signals exhibited contamination and spurious spikes that could not be associated with 

flowfield pressure changes. A better solution would be to use newly available NI Signal conditioner / 

excitation modules that can be remotely controlled by the UDAS and can be located just outside of the 

cam box, very close to transducers. This will minimize the length of cables carrying low voltages and 

minimize possibility of outside effects on transducer signal quality. However, this solution will also be 

somewhat costly. 

NASAlTM-2002-211 894 16 
Kulite calibration. The adopted procedure for Kulite transducers seemed to be satisfactory as 

demonstrated by relatively very good agreements of Kulite DC pressure levels with data from 

conventional static taps (see Fig. 3.5 as an example). As described in Chapter 4, zeros for all Kulite 

transducers were recorded just prior commencement of each test group. These readings, instantaneous 

zero offsets, respectively their differences from the zero offset recorded during the last previous Kulite 

calibration, were used for data correction for transducer zero drift. In future experiments, testing will be 

interrupted every two or three hours to allow for more frequent readouts of Kulite instantaneous zeros to 

further improved the accuracy of DC pressure levels measured with Kulite transducers. 

Once-per-period signal. The OPP pickup worked reliably. The only inconvenience was of the 

current arrangement was that the pickup holder did not allow for position adjustment along the 

circumference of the hexagon cam and therefore the OPP signal could not be aligned with the 'zero' 

blade angle already on the hardware side but it had to be aligned during the data postprocessing (as 

explained in Fig. 3.3). A small design change in the layout of the OPP holder will eliminate this 

inconvenience for the future testing. 

Accelerometer data. The accelerometer arm must be stiffer to avoid arm oscillations in order to 

record blade acceleration more reliably. This is even more true amplitude of blade oscillations larger than 

0.6 dg because a significant increase in blade-tip acceleration. The easiest way will be to shorten the arm 

to a minimum length that will still allow mounting an accelerometer. The accelerometer will not measure 

now the blade tip acceleration but acceleration at a smaller radius. Because the blade is assumed to be 

sufficiently rigid, the instants of zero acceleration and blade-zero angle should be still identical. Smaller 

peak acceleration levels , experienced by the accelerometer will also relieved problems with breaking the 

accelerometer output leads. Further, more sensitive accelerometers can be used. The range of 

accelerometer used and the expected range of measured acceleration levels must be matched better 

than was the case 0 small amplitude oscillations. 

ESCORT/USDA synchronization. Cascade steady data and facility data are acquired by NASA 

GRC labwide system ESCORT. The unsteady data are acquired using a facility dedicated unsteady data 

acquisition system (UDAS). At present, these two systems are completely independable without means of 

mutual communication. All data transfer between the ESCORT and UDAS is done only during the 

postprocessing phase. It would ease the operation of the test facility and simplify the postprocessing 

procedures if ESCORT data, relevant to the cascade steady state operation, were synchronized with the 

data acquisition using UDAS and the these ESCORT experimental data stored concurrently in a suitable 

form in the USDA data files. 

Implication at least some of the above stated suggestions will improve quality and reliability of the 

pressure unsteady data that will be acquired in the follow up phase of the stall flutter experiment for blade 

oscillations at large amplitudes. 

References 

1.1 Buffum, D.H., Capece, V.R., King, AJ ., and EI-Aini, Y.M.: "Oscillating Cascade Aerodynamics at 

Large Mean Incidence," ASME Paper 96-GT -339; also NASA TM-1 07247, 1996. 

1.2 Buffum, D.H. , Capece, VR. , King, AJ., and EI-Aini, Y.M .: "Experimental Investigation of Unsteady 

Flows at Large Incidence Angles in a Linear Oscillating Cascade," AIAA Paper 96-2823; also 

NASA TM-1 07283, 1996. 

1.3 Buffum, D.H. and Fleeter, S.: 'Wind Tunnel Wall Effects in a Linear Oscillating Cascade," NASA 

TM-103690, 1991 . 

1.4 Ott, P., Norryd, M. , and Bolcs, A.: "The Influence of Tailboards on Unsteady Measurements in a 

Linear Cascade," ASME Paper 98-GT -572, 1998. 

1.5 Buffum, D.H. and Fleeter, S.: "Investigation of Oscillating Cascade Aerodynamics by an 

Experimental Influence Coefficient Technique," NASA TM-1 01313, 1988. 

NASAffM-2002-211894 17 
2.1 Boldman, D.R and Buggele, A.E.: "Wind Tunnel Tests of a Blade Subjected to Midchord Torsional 

Oscillations at High Subsonic Stall Flutter Conditions," NASA TM-78998, 1978. 

2.2 Shaw, L.M. , Boldman, D.R., Buggele, A.E., and Buffum, D.H.: "Unsteady Pressure 

Measurements on a Biconvex Airfoil in a Transonic Oscillating Cascade," Journal of Engineering 

for Gas Turbines and Power, Vol. 108, pp. 53-59, 1986. 

2.3 Lepicovsky, J. , McFarland, E.R , Chima, R.V, and Wood, J.R.: "On Flowfield Periodicity in the 

NASA Transonic Flutter Cascade, Part I-Experimental Study," ASME Paper 2000-GT -0572, 

2000. 

2.4 Chima, R.V, McFarland, E.R., Wood, J.R., and Lepicovsky, J.: "On Flowfield Periodicity in the 

NASA Transonic Flutter Cascade, Part II-Numerical Study," ASME Paper 2000-GT -0573, 2000. 

2.5 Lepicovsky, J., McFarland, E.R., Chima, R.V, and Wood, J.R.: "On Flowfield Periodicity in the 

NASA Transonic Flutter Cascade," Journal of Turbomachinery, Vol. 123, pp. 501-509, 2001 . 

3.1 Fronek, D.L. , Setter, R.N., Blumenthal, p.z., and Smally, R.R.: "A Distributed data Acquisition 

System for Aeronautics Test Facilities," NASA TM-88961 , 1987. 

4.1 Lepicovsky, J., Chima, RV., Jett, T.A. , Bencic, T.J ., and Weiland, K.E.: "Investigation of Flow 

Separation in a Transonic-Fan Linear Cascade Using Visualization Methods," International 

Symposium on Flow Visualization, Paper 335, 2000. 

5.1 Gerolymos, G.A. : "Periodicity, Superposition, and 3D Effects in Supersonic Compressor Flutter 

Aerodynamics," International Journal of Turbo and Jet Engines, Vol. 7, pp. 143-152, 1990. 

5.2 Buffum, D.H. and Fleeter, S. : "Oscillating Cascade Aerodynamics by an Experimental Influence 

Coefficient Technique," AIAA Journal of Propulsion and Power, Vol. 6, No. 5, pp. 612-620, 1990. 

5.3 Bolcs, A., Fransson, T.H. and Schlafi, D.: "Aerodynamic Superposition Principle in Vibrating 

Turbine Cascades," AGARO-CPP-468/469, pp. 5-1 - 5-19, 1989. 

5.4 Fransson, T.H.: "Analysis of Experimental Time-Dependent Blade Surface Pressures from an 

Oscillating Turbine Cascade with the Influence-Coefficient Technique," ASME Paper No. 90-GT225, 1990. 

5.5 Smith, S.N.: "Discrete Frequency Sound Generation in Axial Flow Turbomachines," Aeronautical 

Research Council Reports and Memoranda No. 3709, 1973. 

6.1 McFarland, E.R.: "A Rapid Blade-to-Blade Solution for Use in Turbomachinery Design," Journal of 

Engineering for Power, Vol. 106, pp. 376-382, 1984. 

6.2 McFarland, E.R.: "An Integral Equation Solution for Multistage Turbomachinery Design 

Calculations," ASME paper 93-GT -41 , 1993. 

6.3 Chima. RV.: "Explicit Multigrid Algorithm for Quasi-Three-Dimensional Viscous Flows in 

Turbomachinery," AIAA Journal propulsion and Power, Vol. 3, No. 5, pp 397-405, 1987. 

6.4 Chima, R.V.: "Calculation of Tip Clearance Effects in a Transonic Compressor Rotor," ASME 

Paper 96-GT -114, Also NASA TM-1 07216, 1996. 

NASAffM- 2002-211894 18 
Fig. 1.1 Test section of the NASA Transonic Flutter Cascade. 

NASAffM-2002-211894 

AXIAL 

DIRECTION 

+X 

+y 

PITCHWISE DIRECTION 

Fig. 1.2 Airfoil and cascade coordinate system. 

Blade chord, C 89.2 mm 

Leading edge camber angle, v 29.5 dq 

Maximum thickness, t max 0.048 C 

Location of maximum thickness, fP max 0.625 C 

Stagger angle, r 60.0 dg 

Number of blades in the cascade 9 

Blade pitch, S 58.4 mm 

Cascade solidity, CIS 1.53 

Pitching axis, fP ax 0.5 C 

Blade height, h 95.9 mm 

Tab. 1.1 Airfoil and cascade parameters. 

19 
FAR UPSTREAM 

X/C = -1 854 

• • • • 

UPSTREAM 

X/C = -0355 

WALL CORNER 

WALL CORNER DOWNSTREAM 

x/C = + 1.388 

Fig. 2.1. Cascade configuration C_ 20.0 / 24.0. 

,<>'" FAR UPSTREAM = UPSTREAM --?- DOWNSTREAM 

1.0 0.8 W 

a: Ma = 0.80 Ma = 1.00 ::::> 

(f) nTB= 1.058 trTB = 1.097 (f) 

w 

a: 0.8 0.6 11. 

0 a. 

f=: ;: n..-,. 

« -(>" -

(f) 

....J 

f- a. 

0.6 nT

Ma 

B= 

= 

1

0.50 

.022 

~ 

0.4 ....J 

« 

3: ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ 

0 .2 0.6 1.0 0.2 0.6 1.0 0.2 0.6 1.0 

TEST SECTION WIDTH , y / W [ 1 1 

Fig. 2.2. Wall static pressure distributions for cascade configuration C_20.0 / 24. 0. 

NASNTM- 2002-2 11 894 20 
KS1_01 

KS1_03 

KS1_05 

-- KS1 08 - - - - KS1 09 

KS1 10 - - - - KS1 11 

KS1_ 12 

KS1 14 - - I 

Port 1 5.4 mm ( 6.0%) 

Port 2 8.5 mm ( 9.5%) 

Port 3 11.6 mm (13.0%) 

Port 4 14.8 mm (16.6%) 

Port 5 17.9 mm (20.0%) 

FLOW 

DIRECTION 

BLADE KS1 

. __ .. _~ .__._

Instrumented suction surface 

Blade height 95.9 mm 

Height of ports 45.7 mm (from blade root) 

Blade chord 89.4 mm 

Port 6 26.8 mm (30.0%) Port 11 58.0 mm (64.9%) 

Port 7 35.6 mm (39.8%) Port 12 62.4 mm (69.8%) 

Port 8 44.6 mm (49.9%) Port 13 71 .3 mm (79.8%) 

Port 9 49.2 mm (55.0%) Port 14 80.2 mm (89.7%) 

Port 10 53.5 mm (59.8%) Port 15 84.7 mm (94.7%) 

Fig. 2.3 Typical location of pressure ports on the blade suction side. 

KP206 

KP2_01 

KP2_03 

KP2_05 

--KP208 -- -

f 

KP2 07 

Port 1 

Port 2 

Port 3 

Port 4 

Port 5 

- KP209 

KP2 10 - - - - KP2 11 

KP2 12 - -

- 0 - KP2 13 

KP2 14 -L - - KP2_ 15 

5.4 mm (6.1%) 

8.4 mm ( 9.4%) 

11 .7 mm (13.1 %) 

14.8 mm (16.6%) 

17.8 mm (19.9%) 

FLOW 

DIRECTION BLADE KP2 

Instrumented pressure surface 

Blade height 95.8 mm 

Height of ports 45.9 mm (from blade root) 

Blade chord 89.3 mm 

Port 6 26.7 mm (29.9%) Port 11 57.9 mm (64.8%) 

Port 7 35.6 mm (39.9%) Port 12 62.3 mm (69.8%) 

Port 8 44.5 mm (49.8%) Port 13 71.2 mm (79.7%) 

Port 9 49.1 mm (55.0%) Port 14 80.1 mm (89.7%) 

Port 10 53.4 mm (59.8%) Port 15 84.7 mm (94.6%) 

Fig. 2.4 Typical location of pressure ports on the blade pressure side. 

NASAffM-2002-21 1894 21 
EDM ELECTRODE 

NASAffM-2002-2 11 894 

Fig. 2.5 EDM electrode and grooved blade. 

PRESSURE 

SURFACE 

SUCTION 

SURFACE 

Fig. 2.6 Blades prepared for transducer insertion. 

22 
Fig. 2.7 Blades instrumented with Kulite transducers. 

Fig. 2.8 Open cam box routing of transducers leads. 

NASAfI'M-2002-211894 23 
c::> SIGNAL FLOW 

Pressure sensors 

Kulite XCQ-062-15A 

Programming language G 

( LabVIEW) 

PROCESS CONTROL 

Amplifiers Endevco 

4430A 

Computer 

NI PXI-81568 

EXCITATION 

Oscilloscope 

Tektronix 24658 

AID boards 

NI6071E 

Fig. 3.1 Unsteady data acquisition system. 

t 

CHANNEL 

BL2 15 

BL2 14 

BL202 

BL2 01 

ACM_ 1 

BLC15 

BLC14 

BLC02 

BL1 01 

OPP 

TIME 

NASAfTM-2002-2 11894 

BLADE BLADE BLADE 

BL2 BL1 BL2 

0.8 f.ls (1.25 MHz) _ +-+00

1 

E----

26.3 f.ls (38 kHz) 

11.2 f.ls 11.2 f.ls 

24.8f.1s 

26.3 f.ls (38 kHz) 

1.5 f.lS 

) Fig. 3.2 Data acquisition timing. 

24 
I . 

INLET FLOW 

BLADE AT 

'ZERO' ANGLE 

CAM POSITION 

FOR BLADE AT 

'ZERO' ANGLE 

SIX BLADE 

CYCLES 

PER ONE CAM 

REVOLUTION 

MAXIMUM 

VOLTAGE 

30 dg 

31 min 

CAM POSITION 

FOR MAXIMUM 

VOLTAGE 

BLADE 

MOTION 

CAPACITANCE 

\j PICKUP 

DIRECTION 

OF CAM 

CAM ROTATION 

MOTION FOR BLADE 

FLUTTER 

BLADE FREQUENCY 200 Hz 

1.2 r-----i f-----r------r-----~----~---____, 

> 

0.8 

w 0.4 

(') 

« 

I- 0 ---1 

o > -0.4 L.-___ ---' ____ --" ____ ---Io... ____ ........ ____ -' 

10 20 

TIME, 't 

30 

[ms] 

40 

Fig. 3.3 Blade 'zero' angle position in the signal time history. 

NASAffM-2002-211894 2S 

50 
,........., 

> 150 

E 

.......... 

> 

100 

w 

~ 

« r 50 ..J 

0 

> 

200 ,........., o U I I I I I I 

> 0 

::l 00 - -

0 0 0 0 '-- - 0 0 0 0 -

> <1 0 

0 0 - 00 0 0 -

00 

(j) 0 Z 

0 - - r « 0 - - - > 

W 0 

0 0 - -

-200 I I I ...L 1 

20 60 100 20 60 100 

CALIBRATION PRESSURE, P [kPa] 

Fig. 3.4 Typical Kulite calibration results. 

NASAffM-2002-2 11894 26 

I 

L _ _ . 
Blade KP1 07-8eo-00 24-0ct-00 25-Jan-01 20-Mar-01 

Port 01 KP1 01 KP1 02 KP1 03 KP1 04 

slooe mV/kPal 1.248 1.251 1.247 1.247 

IvO mV1 1.35 1.43 1.32 1.28 

s dey mV1 188.8 222.8 245.9 473.7 

Port 02 

slope mV/kPa1 1.612 1.612 1.611 1.609 

Iv. 0 mV1 2.98 3.03 3.03 2.51 

s dey IImV1 294.9 340.5 351.4 650.2 

Port 04 

slope mV/kPa1 1.262 1.265 1.261 1.260 

v. O mV1 -3.41 -3.60 -3.80 -3.98 

s dey mV] 199.8 244.4 250.6 485.4 

Port 05 

slope mV/kPa] 1.477 1.478 1.476 1.475 

Iv_ 0 mV1 5.72 5.50 4.70 4.20 

s dey mV1 274.9 306.4 309.3 582.2 

Port 06 

slope mV/kPa] 1.466 1.470 1.468 1.468 

Iy_ 0 mV] 2.83 2.64 2.72 2.40 

s dey mV] 259.4 295.5 290.3 571.6 

Port 09 

slope mV/kPal 1.460 1.464 

Iy_ 0 mV1 5.65 5.72 

s dey mV1 196.7 231.5 

Port 10 

slope mV/kPal 1.333 1.336 1.335 1.341 

V. O mV] 0.77 0.85 0.60 -0.23 

s dey mV] 214.6 248.0 253.7 510.1 

Port 11 

slope mV/kPa] 1.916 

y. O mV] 3.53 

s dey mV1 857.9 

Port 12 

slope mV/kPal 1.342 1.342 1.344 1.342 

Iv 0 [mY] 0.94 0.48 -2.36 -5.94 

s dey mV1 222.3 256.0 266.1 513.8 

Port 13 

slope mV/kPa1 1.327 1.326 1.323 1.321 

Iv. 0 mV] 2.04 1.86 1.83 1.42 

s dey mV] 210.9 246.6 262.6 509 

Port 15 

slope mV/kPa1 1.372 1.376 1.370 1.369 

Iv. 0 mV] 0.48 0. 14 0.29 -0.08 

s dey mV] 245.4 290.8 275.8 536.6 

Tab. 3.1 Calibration results for blade KP1. 

NASAffM-2002-2 11894 27 
I 

Ion. 

LL 

LL 

ill 

o 

o 

Cf) 

Cf) 

ill 

a: 

n... 

BLADE BL4 - Suction surface 

=0= Bid 51 Static taps (DG_38) 

o 

0.4 

o 

-0.4 

-0.8 

Bid KS1 Kulites 1

51 test (DG_35) 

Bid KS 1 Kulites 2'd test (DG_ 42) 

Bid KS2 Kulites 1

51 test (DG_34) 

Bid KS2 Kulites ~ test (DG_41) 

FLOW INCIDENCE 10 dg 

NO OSCILLATION 

MaIN = 0.50 MaIN = 0.80 

0.2 0.6 1.0 0.2 0.6 

CHORDWISE POSITION, X / C [ 1 ] 

Fig. 3.5 Comparison of static tap data with Kulite average 

pressure data for suction side of blade BL4. 

1.0 

NASAffM-2002-2 11 894 28 

L-. __ . ___ __ ... ___ _~ ... 
Ma,N = 0.50 Data file : C1_35_ 14_KS C 03. hr3 

co 100 

(L 

~ 

Q. 80 

w 

a: 

::) 60 

(j) 

(j) 

w 

a: 40 

(L 200 400 600 800 1000 

TIME, "C [ms] 

co 100 

(L 

~ 

Q. 80 

w 

a: 

::) 60 

(j) 

(j) 

w 

a: 40 

(L 20 40 60 80 100 

TIME, "C [ ms ] 

Fig. 4.1 Uncontaminated raw data signal. 

NASAfIM-2002-211894 29 
MaIN = 0.80 Data file : C1_ 36_ 17_ KS1_ 15. hr3 

CO 100 

(L 

~ 

0. 80 

w 

a: 

::J 60 

Cf) 

Cf) 

w 

a: 40 

(L 200 400 600 800 1000 

TIME, 't [ms] 

Fig. 4.3a Data signal heavily contaminated by 'single point ' spikes. 

MaiN = 0.80 Data file: CC36_ 17_KSC 15. hc3 

ro 100r---------r---------r---------~--------._--------, (L 

~ 

0. 80 

w 

a: 

::J 60 

Cf) 

Cf) 

w 

a: 40~--------~--------~--------~--------~--------~ (L 200 400 600 800 1000 

TIME, 't [ ms] 

Fig. 4.3b Data signal cleaned by the clipping procedure. 

MaIN = 0.97 Data file: CC42_48_KSC01. hx3 

ro 60r---------r---------r---------~--------._--------, (L 

~ 

0. 40 

w 

a: 

::J 20 

(j) 

Cf) 

w a: 0 ~ ________ ~ ________ ~ __________ ~ ________ ~ ________ ~ 

(L 20 40 60 80 100 

TIME, 't [ ms ] 

Fig. 4.2 Bistable flow (pressure jumping) in the transonic region. 

NASAffM-2002-2 11 894 30 

.. ... . ~ 
w 

a: 

~ 60 

en 

en 

w 

MaIN = 0.50 

a: 40----------~--------~--------~--------~------~ 0... 200 400 600 800 1000 

TIME, 't [ ms ] 

ro 100~--------r_--------~--------~----~--~--------~ a... 

.::t:. 

~ 80 ~~~~~~~~~~~~~~~~~~U~~~~~~U 

w 

a: 

~ 60 

en 

en 

w 

a: 0~------~--------~--------~--------~--------~ I 

0... 20 40 60 80 

TIME, 't [ ms] 

Fig. 4.4 Data signal heavily contaminated by temporarily dropouts 

(amplifier saturation). 

MaIN = 0.50 

100 

ro 100~--------~--------~--------~------~--r_------~ 0... 

.::t:. 

~ 80 ~~~~~~~~~~~~~~~~~~U~~~~~~U 

w 

a: 

~ 60 

en 

en 

w 

I 

a: 0~--------~--------~--------~--------~--------~ I 

0... 20 40 60 80 100 

TIME, 't [ ms] 

Fig. 4.5 Data signal cleaned by the clipping procedure. 

NASAlTM-2002-211894 31 
MaIN = 0.50 

C'U 100 

0.. 

.::s::. 

Cl.. 80 

w 

[[ 

~ 60 

(j) 

(j) 

w 

[[ 40 

a.. 20 40 60 80 100 

TIME, 't [ ms ] 

Fig. 4.6 Upper and lower pressure limits. 

MaiN = 0.50 Data file: CC34_ 01_ KS2_ 03. hp3 

C'U 100 

0.. 

.::s::. 

Cl.. 80 

w 

[[ 

~ 60 

(j) 

(j) 

w 

[[ 40 

0.. 20 40 60 80 100 

TIME, 't [ms] 

Fig. 4.7 Data signal cleaned by the patching procedure. 

NASAJTM-2002-2 11 894 32 

I ------ --- -- - ---- ---- -- --- --

-
INLET 

~ [!] 

-===- I 0, REFERENCE AIRFOIL 

Fig. 5.1 Schematic representation of a cascade of airfoils and the numbering 

system used for the influence coefficient method. 

I 0, OSCILLA TlNG AIRFOIL 

Fig. 5.2 Schematic representation of a cascade of airfoils with the reference 

airfoil oscillating (n = 0). 

NASAffM-2002-211894 33 
25 

u.: 

ll.. 

O~ 20 uj~ 

(f)= 

~ uQ. 

0...= 15 

ll.. 

01-

w Z 

o!:!:! 10 

::JO 1-- _ll.. 

Zll.. 

~w 5 

.:x: 0 ~O 

0 

TUNNEL WALL -

-

OSCILLATING +1 AIRFOIL +2 AIRFOIL 

INSTRUMENTED REFERENCE +1 AIRFOIL 

INSTRUMENTED REFERENCE 

-

- TUNNEL WALL 

Fig. 5.3 Configuration to measure the +1 and +2 influence coefficients using two 

instrumented airfoils with the center cascade airfoil oscillating. 

AIRFOIL 

o 0 

CHORDWISE 

DISTANCE, 

xJC [1 ] 

L[) 

Fig. 5.4 Chordwise distribution of the magnitude of the unsteady pressure 

difference influence coefficient. 

25 

20 

15 

10 

5 

o 

AIRFOIL 

NASAfTM-2002-2 11 894 34 

-------- --- ------ - - ----- ---- -
w 10 

0 

z 

W r- 5 0: 

W 

LL 

LL 

0 C:$ 

en C\J -5 

en > 

w C- Real All airfoils 0: 

0... - -10 Imag oscillating 

>- en 

0 a. -1 5 \7 \7 N=5 

<{ 

W 0.. 6. 6. N =2 

f0- g Cf) -20 z 0.2 0.4 0.6 0.8 ::> 

CHORDWISE POSITION, x/C 

Fig. 5.5. Unsteady pressure difference coefficient distribution for 

0 = a dg, Ma = 0.5, and Y = 60 dg. 

w 

0 30 

z 

W r0: 20 

w 

LL 

LL 

0 C:$ 

en C\J 

(f) > 

w co: ........- Real All airfoils 0... -,..-... Imag oscillating 

>- en 

0 a. 

<{ -20 \7 \7 N =5 

W 0.. 

f- a. 6. 6. N =2 

en ........-

z -30 

::> 0.2 0.4 0.6 0.8 

CHORDWISE POSITION, x/C 

Fig. 5.6. Unsteady pressure difference coefficient distribution for 

0 = 180 dg, Ma = 0.5, and y = 60 dg. 

NASAffM-2002-211894 35 

1.0 

[ 1 ] 

1.0 

[ 1 ] 

_ I 
I 

I 

L 

.... ~""""''''' ~ ... .... .... W5-B ........... .. .... Facility ................ .. 

.... 

Fig.6.1a Schematic of the facility with test section shown. 

W5-B 

Facility 

Fig. 6.1 b Schematic of the facility with CFD simulation replacing the test section. 

NASAfTM-2002-2 11 894 36 
z 

;t:-

en 

tv 

o o 

tv 

N 

00 

'0 

W 

-.l 

IV 

c.. 

,:,t. 

...... 

AMPLITUDE 

EFFECTS 

MaIN = 0.5; fBL5 = 500 Hz 

1 .5 , ..... 

ex BL5 

0.6dg 

1.2dg 

2.4 dg 

b 1.0 II 

o 

i= 

« 

0.5 

c 

I 

fJ) 

W 

.. 

, 

, 

, 

\ 

a: 

:J 

fJ) fJ) 

w 

0 1 ' ---j-- I 

0.2 0.6 1.0 

a: 

c.. CHORDWISE POSITION, I C 

Provides largest impact on loading of 

adjacent blades. Also, simulation 

indicates possible wall interference 

at largest amplitude of 2.4 dg. 

[ 1 ] 

---- ---

- ----- ----

MACH NUMBER 

EFFECTS 

ex B

_

5 = 0.6 dg; fBL5 = 500 Hz 

0.5 

MaIN 0.8 

pressure surface 

---- blade BL6 (+1) 

suction surface 

------

blade BL4 (-1) 

......... 

--

........... , ..... 

0.2 0.6 

\ \ 

1.0 

FREQU

EFFE

MaIN = 0.5

0.2 

CHORDWISE POSITION

Provi des significant impact on loading 

of adjacent blades. 

Has li ttle im

adjacen t bl

Fig. 6.3 Summary of CFD predictions. 

- -

._----- .. --- .. --- --- ---- ._----- --
incidence angle 10 dg 

oscillation amplitude 0.0 dq 

oscillation frequency o Hz 

inlet Mach number 0.40 0.50 0.60 0.70 0.80 0.85 

0.90 0.93 0.95 0.97 0.97 0.99 

1.00 1.01 1.03 1.05 1.10 

Tab. 7.1 Test matrix of nonoscillating blade data. 

incidence angle 10 dg 

inlet Mach number 0.5 0.8 1.0 1.1 

oscillation amplitude 0.6 dq 

oscillation frequency 200 Hz 300 Hz 400 Hz 500 Hz 

Tab. 7.2 Test matrix of oscillating blade data. 

Tab.7.3 Blade allocation table. 

NASAlTM-2002-2 11894 38 

'---- - - - - --- --- -- ------
~ - ~ ----- ----- - ------------- ------- -- ----- --~- ---- - 

- --- -- -- ~--- --- ~ 

FLOW INCIDENCE 10 dg 

~ 01 

BL2 BL3 BL4 

TIME RESOL VED SIGNALS (PORT 01) 

100 I I I I 

.. 

60 I20 

CO MaIN = 0.50 C1_36_05_KS1_01. hx3 

0... I I 

.::t:. 100 

0.. 60 

- w 20 a: MaIN = 0.80 CC 36_ 17_KS C 01. hx3 

:::::> 

(f) 100 (f) 

W MaIN = 0.90 CC 36_27_KS C 01. hx3 

a: 

0... 60 

>-

0 

« 

w 20 I-

(f) 

Z 100 :::::> MaIN = 0.93 CC 36_29_KS C 01. hx3 

60 

20 

o 40 80 120 160 200 

TIME, '! [ ms ] 

Fig.7.1a Effects of inlet Mach number on unsteady pressure data. 

NASAffM-2002-211894 39 
C'Cl 

0... 

.Y. 

a. 

w 

0:: 

::J 

(f) 

(f) 

W 

0:: 

0... 

>-

0 

« 

w 

I-

(f) 

Z 

:::::> 

FLOW INCIDENCE 10 dg 

~ 01 

BL2 BL3 BL4 

TIME RESOL VED SIGNALS (PORT 01 ) 

100~--~--~--~--~--~--~--~--~--~--~ 

MaIN = 0.95 

60 

20 

100 

MaIN = 0.95 

60 

20 

100 

MaIN = 0.98 

60 

20 

100 

MaIN = 0.98 

60 

20 

o 40 80 

TIME, 't 

CC 36_30_KS C 01. hx3 

CC 36_33_KS C 01. hx3 

C1_36_35_KS1_01. hx3 

120 

[ms] 

160 200 

Fig. 7.1 b Effects of inlet Mach number on unsteady pressure data. 

NASAfI'M-2002-2 11 894 40 

'- - ---- . .--. -----. -- -.-

I 

I 

I 

_J 
- - - - 

~ 

FLOW INCIDENCE 10 dg 

~ 01 

BL2 BL3 BL4 

TIME RESOL VED SIGNALS (PORT 01) 

100 

MaIN = 1.00 

60 

20 

ctS 

a.. 

.:L 100 

Ma,N = 1.00 

0- 60 

w 20 a: 

::J 

(f) 100 (f) 

W MaIN = 1.01 CC 36_ 46_KS C 01. hx3 

a: 

a.. 60 

>-

0 

« 

w 20 f-

(f) I I I 1 

Z 100 ::J I 

MaiN = 1.02 CC 36_ 44_KSC 01. hx3 

60 ~ -

~ -

20 

.1. .1. .1. 

0 40 80 120 160 200 

TIME, '! [ms] 

Fig. 7.1 c Effects of inlet Mach number on unsteady pressure data. 

NASAffM-2002-211894 41 
l __ ~ 

a... 

W 

0: 

~ 

(f) 

(f) 

W 

0: 

Q.. 

>-

o 

« 

w 

I-

(f) 

Z 

~ 

FLOW INCIDENCE 10 dg 

MaiN: 1.08 

~ 

BL3 BL4 

TIME RESOLVED SIGNALS (PORT 15) 

BL5 

60r---r---~--~--~--~--~--~--~--~---' 

40 

fBL5 : 0 Hz 

60r---r---~--~--~--~--~--~--~--~---' 

50 

40 

fBL5 : 200 Hz 

60r---r---~--~--~--~--~--~--~--~---' 

50 

40 

50 

40 

o 

fBL5 : 300 Hz 

fBL5 : 400 Hz 

20 40 

TIME, 't 

60 

[ms] 

80 

Fig. 7.2 Effects of blade oscillation frequency. 

100 

NASAffM-2002-2 11 894 42 
I -

I 

I 

w 

a: 

::::J 

(f) 

(f) 

W 

a: 

a.. 

>- o 

« 

w 

I--

(f) 

Z 

::::J 

FLOW INCIDENCE 10 dg 

MaIN = 0.80 

~ 

BL3 BL4 

TIME RESOL VED SIGNALS (PORT 15) 

BL5 

80r---r---~--.---~--'---~---r---.---r--~ 

70 

60 

fBL5 = 0 Hz 

80~--r---~--.---~--'---~---r---.--~--~ 

70 

60 

fBLs = 200 Hz 

80~--r---~--~--.---~--'----r---r---r--~ 

o 

fBLs= 500 Hz 

20 40 

TIME, 't 

60 

[ms] 

80 

Fig. 7.3 Effects of blade oscillation frequency. 

100 

NASAffM-2002-211894 43 

---- --- J 
L_ 

ctS 

0.... 

.Y 

Q.. 

w 

a: 

:J 

(j) 

(j) 

w 

a: 

0.... 

>-

0 

« 

w 

I-

(j) 

Z 

:J 

FLOW INCIDENCE 10 dg 

MaIN = 0.50 

~ 

BL3 BL4 

TIME RESOLVED SIGNALS (PORT 15) 

BL5 

100~--~--~--~--~--~--~--~--r---r-~ 

90 

80 

faLs = 0 Hz 

100 

90 

80 

faLs = 200 Hz 

100 

90 

80 

faLs = 400 Hz 

100 

90 

80 

faLs = 500 Hz 

0 20 40 60 80 100 

TIME, 't [ ms ] 

Fig. 7.4 Effects of blade oscillation frequency. 

N AS AfTM-2002-2 I 1894 44 

I 

I 

I 
40 

30 

20 !. 

(\l 

(L 

~ 60 

50 

Q. 

40 

ill 

a: 

:::::> 

CI) 70 CI) 

ill 

a: 

(L 60 

>-

0 

~ 50 

ill 

ICI) 

Z 60 

:::::> 

50 

40 

0 

FLOW INCIDENCE 10 dg 

BL3 BL4 

MaIN = 0.50 

faL5 = 400 Hz 

BL5 

TIME RESOL VED SIGNALS ( SUCTION SIDE ) 

I I I I I I 

Port 1 C C 42_98_KS C 01. hx3 

..... . .... .., 

Port 6 C C 42_98_KS C 06. hx3 

Port 7 CC 42_ 98_ KSC 07. hx3 

Port 15 C C 42_98_KS C 15. hx3 

-

20 40 60 80 100 

TIME, 'T [ms ] 

Fig. 7.5 Effect of position along blade chord. 

NASAfTM-2002-211894 45 
VIEW FROM ABOVE 

VIEW FROM BELOW 

Fig. 7.6 Arm with accelerometers. 

NASAffM-2002-2 11 894 46 

I 

I 

I 

I 

, I 

I 

I 

• I 

I 

I 

I 

I 

I 
faL5 = 200 Hz 

Data file: CC 4C27_ KSCOO. hx3 

CC 4C27_KS2_00. hx3 

1.2 r---~I:__----d~-------.........__--_.., 

> 

w 

C9 

« 

~ o 

> 

0.8 

0.4 

o 

-0.4 L..-_______ ...l--_______ ---I. ___ ---I 

~ 150 r----~---~----r---~---~ 

z 

o 

~ a: 

w 

-l 

W 

o 

o 

o -150 ~---~---~---~---~---~ « 10 

faL5 = 300 Hz 

20 30 40 

TIME, 't [ ms] 

Data file: CC 41_2B_KSCOO. hx3 

CC 4C 2B_KS2_00. hx3 

50 

1.2 r---------r---------,------, 

> 

w 

C9 

« 

~ o 

> 

0.8 

0.4 

o 

-0.4 '----------'------------'------' 

:§ 150 

Z 

o 

f-

« 0 

a: 

w 

-l 

W 

o 

o -150 ~---~----'-----~------'------' « 

NASAffM-2002-211894 

10 20 

TIME, 't 

30 

[ms] 

40 

Fig. 7.7a Once-per-period and accelerometer signals. 

47 

50 
> 

w 

C) 

; 

o 

> 

fBL5 = 400 Hz 

Data file : C1_ 41_29_KS1_00. hx3 

CC 4C 29_KS2_00. hx3 

1.2 ,...-----------,----------r-----, 

0.8 

0.4 

o 

-0.4 '---------....... -----------'-------' 

§ 150 

z 

o 

~ 0: 

W 

---1 

W 

o 

o 

o -150 '-----~---....... ---~-------'~------' et: 10 

fBL5 = 500 Hz 

20 

TIME, 't 

30 

[ ms ] 

40 

Data file: C1_4C 30_KS C OO. hx3 

CC 4C 30_KS2_00. hx3 

50 

1.2 ..--------......---------,------, 

> 

w 

C) 

~ ---1 

o 

> 

0.8 

0.4 

o 

-0.4 1--_______ -'--_______ ---' ___ ____' 

'0) 150 ,...------,------,------r-----r-----, 

Z 

o 

ret: 0 

0: 

W 

---1 

W 

o 

o 0'-----~---~---~--------'---------' et: 10 20 

TIME, 't 

30 

[ms] 

40 

Fig. 7.7b Once-per-period and accelerometer signals. 

AS AfTM-2002-2 I 1894 48 

L-. __ . __ . _____ . ___ ._ . _____ ._ 

50 
Frequency Hz] 200 300 400 500 

Amplitude mm] 0.42 0.42 0.42 0.42 

Velocity I[m/s] 0.265 0.397 0.529 0.661 

Acceleration [m/s2] 166.2 374.0 664.8 1038.8 

Acceleration [gl 16.9 38.1 67.8 105.9 

Tab. 7.4 Expected acceleration paramete~s. 

NASAffM-2002-211894 49 
REPORT DOCUMENTATION PAGE I Form Approved 

OMB No. 0704-01 88 

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gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this 

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 JeHerson 

Davis Highway, Suite t204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503. 

1. AGENCY USE ONLY (Leave blank) 

1

2

. 

REPORT DATE 

1

3

. 

REPORT TYPE AND DATES COVERED 

October 2002 Technical Memorandum 

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS 

Methodology of Blade Unsteady Pressure Measurement in the NASA Transonic 

Flutter Cascade 

VVU-708-28-13-00 

6 . AUTHOR(S) 

J. Lepicovsky, E.R. McFarland, V R. Capece, T.A. Jett, and R.G. Senyitko 

7. PERFORMING ORGANIZATION NA ME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION 

REPORT NUMBER 

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field E-13576 

Cleveland, Ohio 44135 - 3 J 9 J 

9. SPONSORING/MONITORING AGENCY NAME(S) AN D ADDRESS(ES) 10. SPONSORING/MONITORING 

AGENCY REPORT NUMBER 

National Aeronautics and Space Administration 

Washington, DC 20546-0001 NASA TM-2002-21 1894 

11 . SUPPLEMENTARY NOTES 

1. Lepicovsky and R.G. Senyitko, QSS Group, Inc. , Cleveland, Ohio 44135; E.R. McFarland and T.A. Jett, NASA Glenn 

Research Center; and VR. Capece, University of Kentucky, Department of Mechanical Engineering, Paducah, Kentucky 

42002. Responsible person, 1. LepicovsJ...)" organization code 58 10, 2 16---433-1402. 

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE 

Unclassified - Unlimited 

Subject Categories: 07,09 and 02 Distribution: Nonstandard 

Available electTonically at http://gllrs.grc.nasa. goy 

This publication is available from the NASA Center for AeroSpace lnfonnation, 30 1--621-0390. 

13. ABSTRACT (Maximum 200 words) 

In this report the methodology adopted to measure unsteady pressures on blade swfaces in the NASA Transonic Flutter 

Cascade under conditions of simulated blade flutter is described. The previous work done in this cascade reported that the 

oscillating cascade produced waves, which fo r some interblade phase angles reflected off the wind tunnel walls back into 

the cascade, interfered with the cascade unsteady aerodynamics, and contaminated the acquired data. To alleviate the 

problems with data contamination due to the back wall interference, a method of influence coefficients was selected for 

the future unsteady work in this cascade. In this approach only one blade in the cascade is oscillated at a time. The 

majority of the report is concerned with the experimental technique used and the experimental data generated in the 

facility. The report presents a list of all test conditions for the small amplitude of blade oscillations, and shows examples 

of some of the results achieved. The report does not discuss data analysis procedures like ensemble averaging, frequency 

analysis, and unsteady blade loading diagrams reconstmcted using the influence coefficient method. Finally, the report 

presents the lessons learned from this phase of the experimental effort, and suggests the improvements and directions of 

the experimental work for tests to be carried out for large oscillation amplitudes. 

14. SUBJECT TERMS 15. NUMBER OF PAGES 

55 

Transonic flutter cascade; Unsteady pressure measurement 16. PRICE CODE 

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT 

OF REPORT OFTHIS PAGE OF ABSTRACT 

Unclassified Unclassified Unclassified 

NSN 7540-01 -280-5500 Standard Form 298 (Rev . 2-89) 

P rescrib ed by ANSI Std. Z39-18 

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