NASA Technical Memorandum 4806

Flight Demonstration of a

Shock Location Sensor Using

Constant Voltage Hot-Film

Anemometry

Timothy R. Moes

Dryden Flight Research Center

Edwards, California

Garimella R. Sarma and Siva M. Mangalam

Tao of Systems Integration, Inc.

Hampton, Virginia

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

1997

FLIGHT DEMONSTRATION OF A SHOCK LOCATION SENSOR

USING CONSTANT VOLTAGE HOT-FILM ANEMOMETRY

Timothy R. Moes*

NASA Dryden Flight Research Center

Edwards, California

Ganmella R Sarma 1 and Siva M. Mangalam 1:

Tao Systems

Hampton, Virginia

Abstract

Flight tests have demonstrated the

effectiveness of an array of hot-film

sensors using constant voltage

anemometry to determine shock position

on a wing or aircraft surface at transonic

speeds. Flights were conducted at the

NASA Dryden Flight Research Center

using the F-15B aircraft and Flight Test

Fixture (FTF). A modified NACA 0021

airfoil was attached to the side of the FTF,

and its upper surface was instrumented to

correlate shock position with pressure and

hot-film sensors. In the vicinity of the

shock-induced pressure rise, test results

consistently showed the presence of a

minimum voltage in the hot-film

anemometer outputs. Comparing these

results with previous investigations

indicate that hot-film anemometry can

identify the location of the shock-induced

boundary layer separation. The flow

separation occurred slightly forward of the

shock-induced pressure rise for a laminar

boundary layer and slightly aft of the start

of the pressure rise when the boundary

layer was tripped near the airfoil leading

*Aerospace Engineer, 805-258-3054

lResearch Engineer, 757-827-4434

_tResearch Engineer, 757-827-4434

edge. Both minimum mean output and

phase reversal analyses were used to

identify the shock location.

Nomenclature

C

CFD

cp

CVA

FTF

NACA

P

Ps

PSD

r

R 1,R2,R F

airfoil chord length, 10 in.

computational fluid dynamics

pressure coefficient,

(p - ps)/T:i

pressure coefficient

corresponding to local sonic

velocity

constant voltage anemometer

Flight Test Fixture

(NASA Dryden)

hot-film current

National Advisory Committee

on Aeronautics

pressure on the airfoil surface

ambient pressure

power spectral density

dynamic pressure

boundary layer reattachment

location

circuit resistances

This paper is declared a work of the US Government and is not subject to copyright protection in the United States.
S

VDC

X

X/C

o_

AT

Avo

hot-film resistance

boundary layer separation

location

circuit input voltage

circuit output voltage

hot-film voltage

distance from leading edge of

the airfoil, in.

nondimensional chord location

temperature coefficient

of resistance,

approx. 0.004 .Q/_;.,/°C

change from the reference

condition of hot-film

resistance

change from the reference

condition of hot-film

temperature, °C

change from the reference

condition of circuit output

voltage

Introduction

Because of the complex interaction of the

shock with the boundary layer, transonic

shock location and transonic aerodynamic

characteristics in general are highly

dependent on Reynolds number.

Experience has shown it difficult to correctly

simulate the full-scale transonic shock

location in wind tunnels, which has led to

significant errors in predicting aircraft

transonic performance and handling

qualities. 1 Consequently, measurement of

transonic shock locations on a full-scale

aircraft is desired to correlate results with

computational fluid dynamics (CFD) and

wind-tunnel predictions.

This paper presents results from flight tests

that demonstrated the effectiveness of an

array of hot films to sense shock location.

Constant voltage anemometry (OVA) 2-7

was used in conjunction with the hot films in

the first phase of a study to develop a realtime sensor that can locate a steady or

rapidly moving shock. Once developed, this

sensor could provide researchers an

additional tool to improve CFD results and

techniques for transonic wind-tunnel

testing.

Another potential application of this sensor

is to determine shock impingement

location. Shock impingement is also

Reynolds-number dependent because the

shock impinges a boundary layer. This

application includes inlet shock location

sensing for supersonic cruise vehicles.

Information from the sensor could be fed

back in real time to a variable geometry

inlet control system designed to optimize

engine inlet performance.

Traditionally, pressure measurements have

been used to estimate shock location. Hot

films are considered a potential alternative

to pressure measurements for shocklocation sensing for several reasons. These

reasons include easier installation and

operation, higher frequency response, and

better spatial resolution as, in many cases,

the hot films can be placed closer together.

Hot films can also provide more information

on the interaction of the shock with the

boundary layer. This information includes

boundary-layer characteristics such as

turbulence level and separation and

reattachment locations.

Flight tests were conducted using the

NASA Dryden Flight Research

Center's F-15B Flight Test Fixture (FTF)

at an altitude of 20,000 ft and Mach

numbers 0.68 through 0.80. A low-aspectratio wing section using a modified

NACA 0021 airfoil 8 was attached to the

side of the FTF and instrumented with

2
upper surface pressure and hot-film

sensors. The hot-film outputs were

correlated to the shock location estimates

inferred from the pressure measurements.

Both laminar and tripped boundary layers

were investigated. Flight test results and a

proposal for a real-time sensor approach

are presented.

Note that use of trade names or names of

manufacturers in this document does not

constitute an official endorsement of such

products or manufacturers, either

expressed or implied, by the National

Aeronautics and Space Administration.

Experiment Description

Flight Test Fixture

The FTF was used to mount and flight test

this experiment. Installed on the lower

fuselage centerline of the NASA F-15B

aircraft (fig. 1), the FTF uses the same

hardware attachment points as the

standard centerline fuel tank. The FTF was

created as a testbed for generic

aerodynamics research 9 and is 107 in.

long, 32 in. high, and 8 in. wide with a 12 °

elliptical nose section and blunt trailing

edge. A noseboom is mounted at the

leading edge of the FTF to measure pitot

and static pressures and local flow angles.

For this experiment, a total temperature

probe was installed on the left side of the

FTF near the trailing edge. The test article

for this experiment was mounted on the left

side of the FTF (fig. 2), approximately 87 in.

aft of the leading edge and 9.6 in. from the

bottom. Data were telemetered to a ground

control room for real-time monitoring and

recording.

Test Article

The aluminum test article (fig. 3) consisted

of a low-aspect-ratio wing with a modified

EC96-43815-02

Figure 1. The F15-B in flight with the FTF and test article installed.

3
Nose section ____ 0

p!:l_ _ Splitter plate -

Pi tic

Angle.oi.attack

and sideslip vanes

Top view

Left side view

_-Total

/--Suspension \\ temperature /

i,,gs(,) \\p,oba __L

\\ I'°'n

i r107.0 in. -_ 20.4 in. _,1_[ _ 970475

Figure 2. The FTF with the test article.

Figure 3. The test article showing upper surface instrumentation and the grit strip.

NACA 0021 airfoil section. The unswept

wing had an 8.5 in. span and a 10 in. chord.

The lower airfoil surface was flattened and

the trailing edge was blunted to allow

internal installation of pressure ports and

the routing of electrical leads from the

surface-mounted hot films. Sixty-two flush

pressure ports were installed on the upper

surface as shown in figure 4.

Stainless steel tubing with internal diameter

of 0.031 in. was used to plumb the pressure

4
8.50 in.

Leading edge

Hot film

x/c = 0.125 o \

O _ O

Pressure port--_ oO o_ o

\

0 _ 0

x/c=0.250 _ o _ o 20--

o _o21

o _ 022

o _ 023

o _ 024

o _ 025

o -- o26

o -- o27

o -- 028

F Outboard o -- o 29 e,4,,.,_, o _ o 30

o _o31

o _ 032

o _o33

o _o34

x/c=0.525 _ o _ o 35--

O _ O

O _ O

O _ O

O -- O

O _ O

O _ O

O _ O

O _ O

O _ O

x/c=0.875 _ o _ o

Outboard _

Flow

direction

Hot-film

B sensors used in

the analysis

Inboard

edge _

Inboard

10.00 in.

pressure row _ 1.00 in. _ pressure row

Trailing edge J_

I- 4.75 in.

L

970476

Figure 4. Upper surface instrumentation on the test article.

ports. The pressure tubing was routed into

the FTF and connected to flexible tubing

(0.055 in. internal diameter) once inside the

FTF. A polymide sheet containing the hot

films and associated wire leads covered the

entire upper surface of the airfoil. Pressure

port holes were drilled through this sheet.

The actual hot-film sensors were placed at

the same chord locations as the pressures,

as seen in figure 4. The electrical leads

from the hot-film sensors were routed along

the top of the test article outward and

forward and then around the leading edge

and into the test article. The leads were

then directed into the FTF and connected to

the CVA circuits.

Instrumentation

Air Data

The FTF pitot-static measurements were

corrected for position error to obtain

5
freestream measurements of Mach,

ambient pressure, and dynamic pressure.

The FTF total temperature probe was used

with Mach number to obtain freestream

ambient temperature. An angle-of-attack

vane was also installed on the FTF

noseboom and provided local angle of

attack forward of the FTF. The angle of

attack at the test article, however, was

appreciably different from that forward of

the FTF because of fuselage flow

straightening.9 Consequently, no direct

measurements of local angle of attack were

made at the test article. All air data

parameters were measured and recorded

at 52 Hz.

Pressure Measurements

Pressure measurements from the upper

surface were obtained using two 32-port

electronically scanned pressure transducer

modules. These modules measured the

difference between the surface pressure

and a reference pressure. The reference

pressure was from the static pressure

source on the FTF noseboom and

measured using a 0-19 psia 20-bit digital

pressure transducer. The forward port

pressures were measured with a +10 psid

module, and the aft pressure ports were

measured with a +5 psid module.

Approximately 3 ft of pressure tubing were

used between the surface port and the

transducer, and consequently pneumatic

lag was less than 0.02 sec. This situation

was not a problem for these tests because

of the quasi-steady-state nature of the flight

test maneuvers. Pressure data were

measured and recorded at 52 Hz.

Hot-Film Anemometry

The multi-element, hot-film sensor sheet

used in these tests consisted of an array of

45 nickel elements (1 mm long, 0.12 mm

wide, and 0.25 _m thick) on a polymide

substrate. The elements were attached to

13-1_m thick, copper-coated nickel leads

that were routed, as shown in figure 3, to

the airfoil lower surface.

The sensor leads were soldered onto

33-gauge wires that were routed through a

narrow slot on the lower surface into the

airfoil. Coaxial cables connected these

wires to a bank of CVAs inside the FTF.

The unheated resistance of the sensor

elements was a nominal 10 _ on the

ground and dropped to approximately 8.5

at 20,000 ft altitude, where the ambient

temperature was approximately 11 °F.

Appendix A describes the operation of the

hot-film anemometry. The voltage across

the hot film remains constant and thus

is referred to as constant voltage

anemometry. However, the measured

output of the CVA circuit is a voltage that

directly relates to the temperature of the hot

film. The output voltages measured from

the multiple hot-film sensors were presample filtered with a 3-pole Butterworth

filter having a low-pass rolloff frequency of

503 Hz.

The data were telemetered to the ground

station and recorded at 1667 Hz. For each

flight test, only 16 of the hot films

were operational. Between flights the

anemometry system could be reconfigured

to measure a different set of 16 hot-film

sensors. Hot films 20 through 35 were used

for the results in this paper.

Test Approach

Flight data were obtained at an altitude

of approximately 20,000 ft during

stabilized flight, angle-of-attack sweeps,

and acceleration and deceleration

6
maneuvers. Before collecting data, the pilot

reset the current in the hot films and autozeroed the circuit output voltage at

20,000 ft.

To properly configure the system, the reset

and auto-zeroing were done at a Mach

number less than the test Mach numbers

(typically Mach 0.50) and only once in the

flight. The stabilized flight points consisted

of approximately 20 sec of constant altitude

and Mach flight. The angle-of-attack

sweeps were typically +2 ° in amplitude and

were performed at Mach numbers between

0.68 and 0.80.

Acceleration and deceleration maneuvers

were performed at constant altitude

between Mach 0.6 and 0.8 with chord

Reynolds number varied between 1.8 and

2.5 million. The flow was expected to be

laminar before any transonic shock

interaction for the low chord Reynolds

numbers tested.

It was desired to obtain data with both

laminar and turbulent boundary layers

because the shock and boundary layer

interactions depend on the condition of the

boundary layer. 1°-11 For some flights,

grains of 0.02-in. maximum diameter grit

were added at approximately 5 to 8 percent

chord to transition the boundary layer and

therefore obtain data with a tripped

boundary layer upstream of the shock

location.

Results and Discussion

Stabilized flight data were obtained at

20,000 ft altitude and at Mach numbers of

0.68, 0.70, 0.72, 0.74, 0.76, 0.78, and 0.80

with a laminar boundary layer upstream of

the shock. One stabilized test point at

Mach 0.70 was obtained with the grit

strip installed. Hot-film and pressure

measurements showed that the grit strip

made a significant effect; however, it was

unknown whether it created fully turbulent

flow upstream of the shock. Hence, when

discussing results using the grit strip, this

paper refers to the boundary layer as

tripped and not turbulent.

As representative cases, the pressure

distribution and hot-film voltage outputs are

plotted in figure 5 for data at Mach 0.70.

Figure 5(a) shows data for a laminar

boundary layer, and figure 5(b) shows data

for a tripped boundary layer.

From the pressure distributions a difference

between the two rows of pressure

measurements can be seen. The difference

is largest in the region near the large

pressure rise associated with the transonic

shock. The differences are seen at each

tested Mach number and are caused by a

3D flow on the test article. Note that the

boundary layer from the FTF at the test

article is approximately 1.5 in. thick. 9 The

critical pressure coefficient, C_*, also

shown on figure 5, is defined as the

pressure coefficient corresponding to sonic

velocity. 10

For this paper, the shock location is inferred

from the pressure measurements to be

where the shock-induced rapid pressure

rise begins. This definition is consistent with

others in the literature. 1°,12 The location is

somewhat subjective because of the

differences in the inboard and outboard

rows of pressure measurements. For the

laminar boundary layer case shown in

figure 5(a), the rapid pressure rise begins

at approximately x/c = 0.325 (near hot

film 23), whereas for the tripped boundary

layer case shown in figure 5(b) the rapid

pressure rise begins at approximately

x/c = 0.275 (near hot film 21).

7
Cp

Hot film no.

20 25 30 35

'"

-1.4 .................. __ock location ....... [ ................................. 1.................

/

-1.2

-1.0

--.8

--.6

--.4

_.2

0

.2

0

.10

.05

0

-.05

-.10

.1 .2 .3 .4 .5 .6 .7 .8 .9 1.0

x/c

970477

.35

.30

.25

.20

.15 Hot-film

mean

output,

V

Inboard Cp

Outboard Cp

Hot-film output

(a) Laminar boundary layer.

Hot film no.

20 25 30 35

-1.6 I---m I .2 _ Inboard Cp

-1.4 I _/"Pressure-inferred I _ Outboard Cp

v shock location i .1 _ Hot-film output

-,.o -,1

-.8 "_ i[_ _1__/_/_ -.2 Hot-film

mean

Cp Cp* L output,

-.6 - -.3 v

' -'

-.2 ; -.5

.2 P i I _ -.e

' ! ' ' -.7

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0

X/C

970478

(b) Tripped boundary layer.

Figure 5. Pressure and hot-film mean voltage distributions for stabilized flight at Mach 0.7.

8
Minimum Mean Voltage Outputs

Figures 5(a) and 5(b) also show mean

voltage outputs of hot films for Mach

number 0.70. Mean voltages were obtained

by averaging 1 sec of stabilized flight data.

Before the test points, the CVA outputs

were zeroed at Mach 0.50 for the laminar

flow case and at Mach 0.65 for the tripped

flow case. The auto-zero function did not

completely zero the hot films, with as much

as 0.08 V offset remaining. Therefore,

postflight zeroing was done before

analyzing the data. Note the location of the

hot film with the minimum voltage output.

For the laminar boundary layer case shown

in figure 5(a), hot film 21 (x/c = 0.275)

outputs the minimum value. Hot film 21

is forward of the rapid pressure rise

(x/c=0.325) and aft of the minimum

pressure location (x/c = 0.250). For the

tripped boundary layer case shown in

figure 5(b), hot film 23 (x/c = 0.325) outputs

the minimum value. Hot film 23 is aft of the

rapid pressure rise (x/c = 0.275).

Analysis of pressure and hot-film data at

other Mach numbers showed similar

results. Locations of hot-film minimum

mean voltages for laminar boundary layer

cases are correlated with the pressureinferred shock locations as a function of

Mach number in figure 6. At all Mach

numbers, the location of the hot-film

minimum is slightly forward of the pressureinferred shock location.

As mentioned previously, stabilized data for

the tripped boundary layer case were only

obtained at Mach 0.70. Hot-film data were

obtained, however, for a tripped boundary

layer during a level acceleration maneuver.

Only hot films 23, 24, 25, 27, 28, and 30

were operational, and the acceleration was

approximately 3 kn/sec.

Hot film no.

20 25 30 35

.85

I i l i i i l • Hot film with minimum output

Pressure-inferred

shock location

..oi'i"

Mach .75 . _ "

.70 ................... T................... E.............. = ............i....................................... i....................................... i................... i...................

.65

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0

X/C

970479

Figure 6. Comparison of hot-film minimum voltages with pressure-inferred shock locations

for a laminar boundary layer.
Figure 7 shows the Mach number at which

these hot films produced the minimum

value of output during the acceleration.

Figure 7 also shows the pressure-inferred

shock location. At all Mach numbers, the

hot film with the minimum output voltage

occurs aft of the pressure-inferred shock

location. This trend was also seen during a

level deceleration maneuver.

Angle-of-attack sweeps were made to

demonstrate the ability to track the shock

location during a dynamic maneuver.

Sweeps of approximately +2 ° were made at

various Mach numbers.

Figure 8 shows a time history of a quasisteady-state test point followed by an

angle-of-attack sweep at Mach 0.70. For

this case a laminar boundary layer is

forward of the shock. During the first 20 sec

of the maneuver, the angle of attack is held

constant while the Mach number changes

slightly. The angle-of-attack sweep is

shown in the last 20 sec. Only four hot-film

outputs are shown for simplicity.

For the most part, hot film 22 has the

minimum voltage output for the first 20 sec

of the maneuver. As the aircraft decreased

angle of attack, hot film 23 has the minimum

voltage output. As the aircraft pulled up, hot

film 22 has the minimum voltage output,

then hot film 21, and then hot film 20 at the

peak angle of attack. Pushing over to trim

angle of attack, hot film 22 again has the

minimum voltage output.

This movement of the minimum voltage

output is consistent with the movement of

the shock shown in figure 9. Pressure

distributions in figure 9 clearly show the

pressure rise moving aft during the

pushover, moving forward during the pullup,

and then moving back to the original

position.

Freestream

Math

Hot film no.

20 25 30 35

,80

I i I i _1

.78 ......................................................... _....................................... i................... _..............................................................................

.,, ................. 4................... i................... i............ i-,...i................... i................... i................... i................... i..................

iiiioiiiii

.72

iii.iiiiil

i i _oiii

.To...................................... i.................................... i................... i................... i................... i......................................

.68 , , ,

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0

_c

97_

• Hot film with

minimum output

Pressure-inferred

shock location

Figure 7. Comparison of hot-film minimum voltages with pressure-inferred shock locations

for a tripped boundary layer.

10
Hot-film

outputs,

V

Change in

angle of attack

from trim flight,

deg

o.25E _ _

.20

h,2O h_0_" ,._0

.05 = : • ": .. ._2 . _ _LOL

,,i

" " ..... "":' ;" _ '_ _ hf22'

•'_- ;_ _....;"_Lh,_J X-h_," "'._

hf22_ _ / _-hf20

.20 -- hf23 J L_ hf22

.2s , I i- I I I I l

3

2

1

o I

-1

-2

Math

.72

.71

.70

.69

.68

.67 I I I I I I I I I

0 5 10 15 20 25 30 35 40

Time, sec 970481

Figure 8. Time history of stabilized flight and an angle-of-attack sweep at Mach 0.7.

Phase Reversal

During the first 20 sec of the time history in

figure 8, these hot-film outputs are highly

sensitive to Mach-number variations. Hot

films 20 and 21 are in phase with Mach

number (i.e., decreasing in output as Mach

number decreases and increasing in output

as Mach number increases). Hot films 22

and 23 show just the opposite response

and are approximately 180 ° out of phase

with Mach number and are therefore out of

phase with hot films 20 and 21.

As appendix B describes, the phase

reversal between hot films 21 and 22 can

be an indicator of the shock location. We

have hypothesized from figure 8 that the

11
Hot film no.

20 25 30 35

-1,4

_,.2 ............ i.................. i.................. i.................. i.................. .................. ..................

__ Direction _ofincreasing

-1.0 .................. !............. _ _i' angle of attack ................... i..................

_.4

Trim alpha

Trim alpha - 1 °

Trim alpha + 2 °

-.2

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0

x/c

970482

Figure 9. Pressure distributions during an angle-of-attack sweep at Mach 0.7.

shock is near hot films 21 and 22 because

the minimum mean voltage output varies

between these hot films for the first 20 sec

of the maneuver. The phase reversal

concept also indicates this hypothesis

because hot films 21 and 22 are out of

phase with each other.

During the angle-of-attack sweep, we can

see that hot films 20 and 23 are basically

out of phase with each other the entire

time, indicating that the shock-induced

phenomena is between them. Another

interesting feature from this low-frequency

analysis is that hot film 22 exhibits a

frequency approximately twice that of the

other hot films during the angle-of-attack

sweep. This feature indicates that the shock

oscillates around hot film 22 during this

maneuver.

Spectral analysis can identify phase

reversal. As stated previously, the hot-film

data were pre-sample, low-pass filtered

with a rolloff frequency of 503 Hz and

recorded at 1667 Hz. A 6-sec portion of the

constant angle-of-attack data from figure 8

was analyzed using Fourier transforms with

resulting phase angles from hot films 20

through 23 shown in figure 10.

At low frequencies clearly hot films 20

and 21 are roughly 180 ° out of phase from

hot films 22 and 23. At higher frequencies

the differences in phase angles are not

distinguishable because of spectral noise.

Phase reversal about the hot film with the

minimum output was seen using this

technique at all Mach numbers tested.

However, additional low-pass filtering of the

signals was required at the higher Mach

numbers before using the Fourier transform

phase angle analysis.

12
150 :, ,; ' ::i:" ,I :": __.

I I II • I': '" I I * h Ii I d

I : "'* Ji ;_ II I III ** l_ l ;

l Hot films 22 and 23 ,_.. :1 , H '_ :,-':: ., , t : _:. ,: :

100 .._:;_,,_ : _g:'_i_: L, rJ;,:tl,.'__, l,_,J1't:,_:_:;:,::

:, ._ t , , , _. : :, :: :: : iL., _; !.E _ _ i_l_ _ _l.: :.. ,:_ i:: :,

,,: " ' "!f, " " ._ ", '-, , ' ,:

,. ,_: t ,:.,_ _ _ !,, _,'

5o ' _ _ti :_i_ "_i! ,!I! ,_:.,

, ;l':_ _; " lift:: .;,-1I/llll[lii:i_ll! i_" i' ":":I ':II'

, , , : ,, l, _!_|if , , ,," I! _:, , , ,,,i ;, , , :

,: ,:. , ! :, ,*, . , ,

angle, 0 :: :: ,: :: ,3 !: ,, _, .I

,,, ,

: : i .... *_'J , *

deg I ::.. •: ,,.j 4.,.,,,:, ,,_...... . :: :

•.. _ . ::_ .... ,:, , :_ , , • " , : -,' " _: :, ,", ,.'i. : :

,! .... :, :::: h 'F!;_I: , : : ::_ .

:" ':'I _:"=: "' ' : ' " •

I- _. • • l I l l .** * • "l .l : :.:, • :!!::!.: , , . il ,. -.. -50 : :,I ::::::': '_ ' ' f i: ' 'if:

I Hot filme 20 and 21 :_'; ",L ,: ' :: ' i " _-"i i, , : "_: h :,! "E

. ' :,.• 1,1 _.,t ,, 0.,:I . ,_:_,

- 100 _ ' :" ' "!i':

- __i y " 1__? ..,1Ii__i' _!!It.,',: ii' i_i_!_ ll,i_-l_!!liti

t : , ,1t ,.

I I I II , I I" * "

-150 i : _ _

' ' / _ ::

i i I ]1 I - _:il

0 5 10 15 20 25 30

Frequency, Hz 970483

Figure 10. Phase angles from hot films 20 through 23 during a Mach 0.7 quasi-steady-state

test point.

Flow Structure

These flight tests provide an initial look into

transonic shock location detection using hot

films with CVA. The results show that both

minimum mean voltage and phase reversal

are potential indicators of the shock

location. The hot films, of course, are

surface measurements that identify

features in the complex off-surface

interactions between the shock and the

boundary layer.

Numerous authors have described shock

and boundary layer interactions in

detail. 1°,13,14 Typically, the adverse

pressure gradient induced by the shock

would cause local separation of

the boundary-layer flow. The location of

the separation is related to the condition of

the boundary layer. As discussed in

appendix A, the hot films in the present

study were operated at current levels

sufficient to measure velocity fluctuations

over the surface and therefore to identify

separated flow regions.

Previous hot-film experiments 4 have shown

that flow separation in the boundary would

lead to a minimum output in a hot-film array

because of reduced convective heat

transfer. Consequently, it was hypothesized

13
that the primary factor creating a local

minimum in the hot-film array voltages was

boundary layer flow separation mduced by

the adverse pressure gradient associated

with the shock. The flow separal,on begins

near the location of the hot f,tm _sth the

minimum output.

At some transonic Mach n_,-"_,e,s two

distinct minimums occur ,n 1_e hot-film

outputs as seen in figure 5 T,_,__second

minimum is suspected to _n3,ca,.eflow

reattachment. Both sepa',_'.,_n and

reattachment points would be ,n re_,ons of

reversed flows and hence a stagnation

point would occur and convecl,ve heat

transfer would be diminished ThIs srtuation

would cause the hot-films to heat up and

the output voltage to drop at these

locations.

A comparison of hot-film output minimums

with the pressure distributions in figure 5

indicates that the separation region began

farther upstream of the shock when the flow

was laminar than when it was tripped. This

result is consistent with the fact that

turbulent boundary layers are more difficult

to separate.

Another observation from figure 5 is that the

hypothesized separation and reattachment

are farther apart for the laminar boundary

layer case. The separation and

reattachment points are separated by

12.5-percent chord for the laminar flow

case and only 7.5-percent of the chord for

the tripped flow case. Figure 11 graphically

displays the shock and boundary layer

features that have been hypothesized from

the hot film and pressure distributions.

Further evidence of this separated flow

region can be seen in the power spectral

densities (PSDs) of the hot films. PSDs of

three normalized hot-film signals and

normalized angle of attack are shown in

figure 12 for data obtained with a laminar

boundary layer at Mach 0.70. The signals

were normalized by dividing the signal by

its standard deviation before obtaining the

PSD. Hot-film minimums were measured at

hot films 21 and 26, indicating a separated

flow region between them (fig. 5(a)). PSD

analysis of the noseboom angle-of-attack

signal showed a peak near 13 Hz at this

Mach number because of an unknown

source.

Flow [_

bubble

Boundary layer

T 970496

(a) Laminar flow upstream of the shock.

Figure 11. Hypothesized flow structure near the shock.

14
bubble L

Boundary

layer

----f--

970484

(b) Tripped flow upstream of the shock.

Figure 11. Concluded.

Figure 12 shows PSDs from hot films 19,

25, and 27. A dominant flow feature is at

13 Hz seen in hot films 19 and 27 but not in

hot film 25. The lack of the 13 Hz signal in

hot film 25 supports the hypothesis that the

boundary layer was separated in this region

102 =

101

100

Normalized 10-1 PSD

10-2

10-3

Stronger roll-off

of hot film 25

Hot film 25

E _ Hot film 27 --/ .__---" . /

-- .o

- X.J '

I I I I I llll I I I I

101

Frequency, Hz

10-4 IIIII I I I I IIIII

100 102 103

97O485

Figure 12. Power spectral densities of hot films 19, 25, and 27 and angle of attack during a

stabilized test point at Mach 0.7.

15
and thus prevented the 13-Hz phenomena

from impacting surface hot film 25. At

higher frequencies the magnitude of the

PSD for hot film 25 drops off more rapidly

than the other signals.Tests at other Mach

numbers also showedthat this observation

was typical of PSD magnitudes from hot

films between two minimums.

Real-Time Sensor

One objective of this phase of the study was

to obtain an approach for a sensor for realtime shock location. A potential approach is

to identify the shock location based on the

minimum mean voltage from the array of

CVA outputs. As previously discussed and

shown in figure 5, the hot film with the

minimum mean output is near the rapid

pressure rise caused by the transonic

shock. A second minimum is sometimes

evident slightly downstream of the first.

These two minimums were hypothesized to

be the point of boundary layer flow

separation induced by the shock and the

reattachment point. In some cases, the hot

film at the reattachment point outputs a

lower voltage than the hot film at the

separation point. Therefore, any algorithm

using this approach would need to

differentiate between these two points.

A shock location algorithm was written to

analyze acceleration and angle-of-attack

sweep maneuvers using the minimum

mean voltage output approach. In both

cases the boundary layer was laminar

before the transonic shock. The shock

location algorithm obtained hot-film

mean voltages over 0.01-sec intervals

to simulate a 100-Hz sensor output.

Figure 13 shows results from analyzing hot

films 20 through 35 during a level

acceleration maneuver. Figure 13(a) shows

.82 i

.80

.78

.76

Mech

.74

.72

.70

.68

0 5 10 15 20 25

Time, sec 970486

(a) Mach-number time history.

Figure 13. Time history of a level acceleration maneuver demonstrating the shock location

algorithm output for a laminar flow boundary layer.

16
.55 i 32

x/c

.5O

.45

i ki

t i

Pressure-inferred I i

shock locatlon_ I ,&-

IFI 1 I, F

.35 "'

i ," - i _ Minimum mean voltage

ouuocton .30, _" t

.25

.20

0 5 10 15 20

Time, soc

A.--=_

Si

• i 30

28

26

24

22

2O

18

25

970487

Hot film

no.

(b) Comparison of locations of the minimum mean voltage output with the pressure-inferred

shock.

.55 32

___.& .50 ........................................ _.............................................................................................................. _.............................................. 30

Pressure-inferred t .

shock Iocation-_ i .... 4["

\ J I

.45 28

," [-

.40 26

i .." i Hot film

x/¢ no.

.35 24

,dk.i•••.. &•""" Shock location i r---j _ _ algorithm output

D3O 4 22

J

.25 i , 20

.20 _ 18

0 5 10 15 20 25

Time, sec 970488

(c) Comparison of the shock location algorithm output with the pressure-inferred shock

location.

Figure 13. Concluded.

17
the Mach time history. Figure 13(b) plots

the hot film that outputs the minimum value.

Figure 13(c) is a result of the shock location

algorithm that chooses the forward hot

film when there are two distinct minimums

(as in figs. 5(a) and 5(b)).

The location of the shock inferred from the

pressure measurements is also shown in

figures 13(b) and 13(c). As can be seen,

the shock location algorithm output

consistently locates the minimum one or

two hot films ahead of the inferred shock

location.

The angle-of-attack sweep plotted in

figure8 was also analyzed with and

without the shock location algorithm, and

figures 14(a), 14(b), and 14(c) show the

results. Without the shock location

algorithm, there is a part of the maneuver

where the minimum location oscillates

between hot films 22 and 26 and stabilizes

at hot film 26. The pressure-inferred shock

location, however, is always before hot

film 26. With the shock location algorithm,

the chosen hot film is always slightly

forward of the inferred shock location, as

expected for the laminar flow case, and the

algorithm adequately tracks the shock.

Spectral analysis, and in particular phase

reversal, could also be used in real-time

shock location identification (although for

this study, no phase reversal algorithm

results are presented). Although phase

reversal can indicate the location of the

shock, such reversal is not a necessary

condition because phase reversal requires

Change in

angle of attack

from trim flight,

deg

2.5

2.0

1.5

1.0

.5

-.5

- 1.0

- 1.5

0 5 10 15 20

Time, sec

(a) Angle-of-attack time history.

25

970489

Figure 14. Time history of an angle-of-attack sweep demonstrating the shock location

algorithm output for a laminar flow boundary layer.

18
.400 i 26

ressure-inferred

•

350 _ i i 24

x/c .325, 23 Hot film no.

.300 22

.275 21

.250 20

.225 19

0 5 10 15 20 25

Time, sec 970490

(b) Comparison of locations of the minimum mean voltage output with the pressure-inferred

shock.

.425 27

.400 .............................................................................. i............................................................................... !....................................... 26

Pressure-inferred i i

.350 ..................................... _I :......................... ___ [.................................................................................................................... 24

/ \ i

x/c .325' _-a,-4,-a_ ..... \ i i

• -'_-_-_ ................... _"-J .'-&-"A--i-'A-_i-dh_A--A ..... 23 Hot film

i \ / i no.

.3oo i _ 22

,275 --Shock location / 21

algorithm output-J i

.250

i = 20

.225 i ! 19

0 5 10 15 20 25

Time, sec 970491

(c) Comparison of the shock location algorithm output with the pressure-inferred shock

location.

Figure 14. Concluded.

19
oscillatory motion of the shock at a

frequency higher than the desired shock

location sensor output frequency.

Concluding Remarks

Flight test data were obtained from a

shock location sensor comprised of hot

films controlled with constant voltage

anemometry. Tests were flown on the

F-15B Flight Test Fixture (FTF) at NASA

Dryden Flight Research Center using a

modified NACA 0021 airfoil attached to the

side of the FTF. Hot-film and pressure

measurements were obtained along the

chord of the test article.

The shock location inferred by pressure

measurements was correlated with the

hot-film outputs. For laminar flow upstream

of the shock, the array of hot films showed

a minimum output slightly upstream of the

pressure-inferred shock location. For test

conditions with tripped (partially turbulent)

flow upstream of the shock, the array

of hot films showed a minimum slightly

downstream of the pressure-inferred

shock location. The hot-film minimum is

hypothesized to correspond to the start of a

shock-induced separation bubble inside the

boundary layer.

A shock location algorithm simulated

100-Hz output of a real-time sensor.

Results from the algorithm were presented,

showing consistent movement of the

hot-film array minimum with the shock.

A phase-reversal analysis was also

presented in the paper. Phase reversal is a

technique that can be used to locate the

hot-film array minimum if the location is

oscillatory.

20
Appendix A

Use of CVA for Shock Location Sensing

The interaction of a shock with a boundary

layer increases ambient air temperature

and changes the flow velocity along a

surface. Surface-mounted hot films could

potentially sense both of these changes.

Figure A1 shows the basic CVA circuit used

to control the hot films. A complete analysis

of this circuit has been presented. 2

Note that whereas the hot film is maintained

at a constant voltage (Vs), the circuit output

voltage (V o) does change because of hotfilm resistance changes. The output voltage

change from a reference condition is

related to the resistance change as

ARs

= t%-j

For this application, the reference condition

refers to a low-speed flight condition at the

test altitude in which no shock was incident

on the sensor. The fractional change in the

hot-film resistance is related to the change

in the temperature in the hot film by

ARs (x AT

R s ! +o_ AT

where o_ is the temperature coefficient of

resistance of the nickel hot-film material,

and AT is the change in temperature from

the reference condition. Therefore, the

output voltage change due to temperature

change is

hV o = -R 21sll +_ ATAT)(z

All hot films in the array were made from

the same sheet of nickel film, and therefore

they had the same temperature coefficient

of resistance o_ of approximately 0.004 £Z/

_2/°C. R 2 is a circuit constant that is a

precision resistor (+0.1 percent tolerance).

Is varies as the hot-film temperature

changes; however, it was initially set at the

low-speed reference condition such that the

sensitivity factor, R21s, was identical for

each hot film in the array at the reference

condition.

Figure A2 shows a schematic of the

operation of the CVA system. This

R1

v°c°--I I

RF Vs R2

t 0 Vo

Figure AI. Basic CVA circuit (U.S. Patent 5074147).

970492

21
arrangement was designed to obtain highly

discernible voltage outputs from the hot-film

temperature changes. At the stabilized lowspeed reference flight condition,the aircraft

pilot toggled the CVA Reset switch, which

started the Auto Current Control sequence.

This sequence set the hot-film current

levels such that the sensitivity factor R21s

would be identical in each hot-film CVA

circuit.

After completion of the Auto Current

Control sequence, the pilot toggled the

Auto Zero switch. This configured the Auto

Zero unit to subtract the static output

voltage (V o) measured at the reference

condition from the CVA outputs. This step

was necessary because the static output

voltage (Vo) could be as large as 15 V for a

100-mA current through the hot film. The

voltage changes resulting from the shock,

however, were only tenths of volts. To

discern these changes with higher

resolution, the static Vo measured at the

reference condition was removed from the

measured output.

The CVA has been shown to be sensitive

primarily to ambient temperature changes

at low-current (low-overheat) levels and

sensitive primarily to velocity fluctuations at

high-current (high-overheat) levels. 7 Flight

tests at low-current levels (I s - 30 mA) were

largely unsuccessful in producing CVA

outputs that correlated well with the

pressure-inferred shock locations. Flight

tests at high-current levels (I s - 100 mA),

however, produced CVA outputs that did

correlate well with the pressure-inferred

shock location.

This result suggests that the primary effect

of the shock on the hot films was to change

the convective heat transfer caused by the

velocity fluctuations in the boundary layer.

Consequently, high-current levels were

used for the results presented in this paper.

CVA reset

.witch

Auto current I

i°ntr°ivSCVAV° 'I_ "" - ' AutoierOAuto.WitChzero

-_Output

970493

Figure A2. CVA and Auto-Zero arrangement.

22
Appendix B

Phase Reversal Illustration

Phase reversal between hot films can be

used as an indicator of shock location. The

phase-reversal phenomena is explained

using figure B1 (a), which graphs

hypothetical voltage outputs from eight hotfilm sensors at five time points. In this

figure, we assume that the shock location is

identified by the minimum voltage output. At

the initial time, the shock is located at

sensor $4. As time changes, the shock

moves between sensors $5, $4, and $3 as

indicated by the black dot.

Figure Bl(b) shows the time histories for

each of the sensors. As can be seen,

sensors $1, $2, and $3 are in phase with

each other. Sensors $5, $6, $7, and $8 are

also in phase with each other. However,

these two groups of sensors are completely

out of phase with each other. Furthermore,

sensor $4 varies at twice the frequency of

the other sensors.

Hence, if the shock is oscillating about a

particular hot film, we would expect to see a

phase reversal between hot films on either

side of that particular hot film. We would

also expect to see the particular hot film

about which the shock is oscillating to

exhibit a frequency twice the oscillation

frequency.

23
o Sls2_;3V s5_ _"_

Volts //"

11

S1 $2 $3 S4 S = _ ._" $8

Ol__/__i_ ' _'f Volts - 1 -- _,

-2 t2

Sl S2 $3 _4 $5 S6 $7 $8

vo.,-__ _

- 2 t3

0

Volts -i t

s,,s__ s, s, s_s_

-. _ T,. =_ f t4

V

Sl S2 S3 _4 S5 $6 $7 S8

0 I I I I" i_

970494

B1 (a) Output voltage as a function of sensor

at five points in time.

tl t2 t3 t4 t5

E

i lP

i i

Sl

S2

S3

S4

S5

S6

$7

$8

970495

B1 (b) Output voltage as a function of time at

eight sensor locations.

Figure B1. Phase reversal illustration.

24
References

1Loving, Donald L., Wind-TunnelmFlight

Correlation of Shock-Induced Separated

Flow, NASA TN D-3580, Sept. 1966.

2Sarma, Garimella R., "Analysis of a

Constant Voltage Anemometer Circuit,"

IEEE Instrumentation and Measurement

Technology Conference, Irvine, California,

May 1993.

3Mangalam, S. M., G. R. Sarma, S. Kuppa,

and L. R. Kubendran, "A New Approach

to High-Speed Flow Measurements

Using Constant Voltage Anemometry,"

AIAA-92-3957, July 1992.

4Mangalam, S. M., G. R. Sarma, and S.

Kuppa, "Quantitative Flow Diagnostics

Techniques for Unsteady Aerodynamics,"

2nd Pacific International Conference on

Aerospace Science & Technology,

Melbourne, Australia, Mar. 1995.

5Harvey, Don, Subrahmanyam Kuppa,

Garimella R. Sarma, and Siva M.

Mangalam, "Demonstration of the

Constant Voltage Anemometer (CVA) in

the National Transonic Facility (NTF),"

ASME, Aug. 1995.

6Harvey, Don, S. Kuppa, Daryl Sinclair, and

Kurt Lotter, "Demonstration of an Advanced

Flow Diagnostics System in the AEDC 16T,"

AIAA-95-2833, July 1995.

7Kegerise, Michael A. and Eric F. Spina, "A

Comparative Study of Constant-Voltage

and Constant-Temperature Hot-Wire

Anemometers in Supersonic Flow," 3rd

International Symposium on Thermal

Anemometry, July 1996.

8Abbot, Ira H. and Albert E. Von Doenhoff,

Theory of Wing Sections, Dover

Publications, Inc., New York, 1958.

9Richwine, David M., F-15B/Flight Test

Fixture I1: A Test Bed for Flight Research,

NASA TM-4782, Dec. 1996.

1°Shapiro, Ascher H., The Dynamics and

Thermodynamics of Compressible Fluid

Flow, The Roland Press Company, New

York, 1954.

11Schlichting, Hermann, Boundary-Layer

Theory, seventh edition, McGraw-Hill Book

Company, 1979.

12pearcey, H. H. and D. W. Holder,

"Examples of the Effects of Shock-Induced

Boundary-Layer Separation in Transonic

Flight," Ministry of Technology, Aeronautical

Research Council, R. & M. No. 3510,

Jan. 1954.

13Seddon, J. and E. L. Goldsmith, Intake

Aerodynamics, AIAA Education Series,

AIAA and BSP Professional Books, 1985.

14Delery, J. and J. G. Marvin, "ShockWave Boundary Layer Interactions:'

AGARDograph No. 280, Feb. 1986.

25
REPORT DOCUMENTATION PAGE Approved OMB No. 0704-01 88

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, 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 Re )orts. 1215 Jefferson Davis Highway, Suite 1204. Arlington_

VA22202-4302, andto the Office ofManagementand Budget,PaperworkReductionProject(0704-0188),Washington,DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORTTYPE AND DATES COVERED

August 1997 Technical Memorandum

4.TITLE AND SUBTITLE

Flight Demonstration of a Shock Location Sensor Using Constant Voltage

Hot-Film Anemometry

6. AUTHOR(S)

Timothy R. Moes, Garimella R. Sarma, and Siva M. Mangalam

7.PERFORMINGORGANIZATIONNAME(S) ANDADDRESS(ES)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

5. FUNDING NUMBERS

WU 529 21 34003800F15

8. PERFORMING ORGANBATION

REPORT NUMBER

H-2191

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

TM-4806

11. SUPPLEMENTARY NOTES

Presented at the 1997 Society of Flight Test Engineers Symposium, Orlando, Florida, August 18-22, 1997. Timothy R. Moes, Dryden

Flight Research Center, Edwards, California; Garimella R. Sarma, and Siva M. Mangalam, Tao of Systems Integration, Inc.. Hampton,

Virginia.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified--Unlimited

Subject Category 06

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

Flight tests have demonstrated the effectiveness of an array of hot-film sensors using constant voltage

anemometry to determine shock position on a wing or aircraft surface at transonic speeds. Flights were

conducted at the NASA Dryden Flight Research Center using the F-15B aircraft and Flight Test Fixture (F'TF).

A modified NACA 0021 airfoil was attached to the side of the FIT, and its upper surface was instrumented to

correlate shock position with pressure and hot-film sensors. In the vicinity of the shock-induced pressure rise,

test results consistently showed the presence of a minimum voltage in the hot-film anemometer outputs.

Comparing these results with previous investigations indicate that hot-film anemometry can identify the location

of the shock-induced boundary layer separation. The flow separation occurred slightly forward of the shockinduced pressure rise for a laminar boundary layer and slightly aft of the start of the pressure rise when the

boundary layer was tripped near the airfoil leading edge. Both minimum mean output and phase reversal

analyses were used to identify the shock location.

14. SUBJECT TERMS

Transonic Shocks, Shock Impingement, Boundary Layers, Hot-Films,

Anemometry, Constant Voltage Anemometry

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION

OF REPORT OF THIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified

NSN 7540-01-280-5500 Available from the NASA Center for AeroSpace Information, 800 Elkridge Landing Road,

Linthicum Heights, MD 21090; (301)621-0390

15. NUMBER OF PAGES

28

16. PRICE CODE

A03

20.LIMITATIONOFABSTRACT

Unlimited

Standard Form 298 (Rev. 2-89)

Prescnbed by ANSI Sld, Z39-18

298-102