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NASA TECHNICAL MEMORANDUM NASA TM-88385

(NASA-TM-88385) 2N EXFIfiJHENTAL N86-24936

INVESTIGATION OB IOCATION OF BOUNDAR Y LAX£B

TRANSITION ON THE 8ACA 0012 GSIHG SOBPACI

HOT FILM GAGES (National Aeronautics and Unclas

Space AdfliiDistratioc) 10 p HC A02/MF A01 G3/34 U3118

AN EXPERIMENTAL INVESTIGATION ON LOCATION OF BOUNDARY LAYER

TRANSITION ON THE NACA 0012 USING SURFACE HOT FILM GAGES

Wang Tie-cheng

Translation of Acta Aeronautica et Astronautica Sinica, Vol. 5

No. 4, Dec. 1984, pages 401-405.

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

WASHINGTON D.C. 20546 MARCH 1986
AN EXPERIMENTAL INVESTIGATION ON

LOCATION OF BOUNDARY LAYER TRANSITION

ON THE NACA 0012

USING SURFACE HOT FILM GAGES

"WANG Tie- cheng

Nanjing Aeronautical Institute

ABSTRACT

Location of the boundary layer transitions was investigated

by attaching the M-1 type surface hot film gages, which were /401*

developed on our own, flush to the upper surface of a NACA 0012

airfoil. As shown in a diagram, the location of the boundary layer

transitions changed as the angle of attack of the airfoil was

varied. Experimental results show that the location of boundary

layer transitions using hot film gages is a feasible and practical

technique.

1. INTRODUCTION

It is extremely important to determine the location of the

interim (transition) region from laminar to turbulent boundary

layers on a solid wall. Determination method using surface hot film

gages is one of the techniques developed for this purpose. The

advantages of this technique are small disturbance on flow (as

compared with projection method), convenience for measurements

and being insensitive to noise and pressure gradients.

Investigation on the location of the boundary layer transitions

*Numbers in margin indicate foreign pagination
using surface hot film gages has been reported in literature L1 ]

and£2]. The characteristics of the experiment described in this

paper are attachment of several surface hot films flush to the

upper surface of the airfoil to locate the boundary layer

transitions and simultaneous use of three different display units

to show the signals involved in these films. The objective of

this design is to locate the boundary layer transitions at

different angles of attack through comparison and verification.

2. PRINCIPLES

Both the projection and laser methods for locating the

boundary layer transitions are based on the big differences in

the speed types and the turbulence level between the laminar and

turbulent boundary layers. In contrast, with the hot film method,

the location of these transitions is measured based on the

different characteristics of the changes in the wall shear stress

at the laminar and turbulent boundary layers, that is, at the same

Reynolds number, the wall shear stress at the laminar flow region

(equivalent to the time average at turbulent flow) is much smaller

than at the turbulent flow, the root mean square of the shear

stress is smaller in the former region than in the latter, and the

pulse wave patterns in these two different regions are also quite

different. After the hot film gages on the model surface are connected to a OTA circuit (constant temperature hot-film and hotwire anemometer), the three quantities (relative values) mentioned

above can be measured by using different display units to determine

the location of the transitions. Details about these measurements

can be found in reference [3J .

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In experiments determining the location of the boundary

layer transitions, only the relative changes in the wall shear

stress need to be measured. Hence, the technique described inC1]

and [2] can be used where the voltage output (time average), E is

directly used to express the wall shear stress (time average),

the root mean square of the pulse voltage, E for that of the /402

jTins

pulse wall shear stress, and the voltage oscillogram for the pulse

wave pattern of the wall shear stress. Here, Both E and E have

rms

been corrected by subtracting their individual initial readings.

3. EXPERIMENTAL

The experiment was carried out in a 2-D low turbulence level

wind tunnel at the Nanjing Aeronautical Institute in 1983. The test

section of the tunnel is 6m long, 0.3m wide and 1.2m high with the

highest wind speed of 42 m/s and the turbulence level less than

0.08%. The chord length of the NACA 0012 airfoil model, b is 0.15m

and its extended length, 0.3m. The model spanned two side walls in

the test section of the tunnel. Assuming x being the distance from

the front edge along the chord length, each M-1 type hot film was

attached flush on the upper surface of the airfoil at x/b=0.4, 0.5,

0.6, 0.7 and 0.8. In the direction of flow, the hot film (thermal

film) is 0.05mm long by 0.005mm thick, the thickness of the film

base being 0.03mm. Grooves were prepared in advance on the model's

surface on which the hot films were to be attached. No steps

between the film base and the model's surface were observed after

the attachment.

A tsi 1050 type hot-wire, hot-film anemometer was used in this

test, together with three displays, tsi 1050 numerical voltmeter

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(measuring the mean, E), tsi 1070 type root mean square voltmeter

(measuring E _ ) and NS-5511 memory oscilloscope (giving the pulse

wave patterns).

4. EXPERIMENTAL RESULTS

The experimental results are shown in Figs. 1 and 2. Fig. 1

shows the E and E vs o(, angle of attack curves with different rms

x/b. Fig. 2 shows the oscillograms at five typical angles of attack with x/b=0.7. The vertical deflection sensitivity and /403

scanning speed of the oscilloscope were kept constant during the

test. Each division along the horizontal line represents" 2ms.

The similar wave patterns were found with other x/b values. The

Reynolds number

NACA-type 0012 airfoil is a typical airfoil with medium

thickness. The largest thickness appears at J>0% length of the airfoil. It can be considered that within the range of angle of attack

investigated, the film-attached surfaces are in the adverse pressure

gradient region. The larger the angle of attack, the larger this

gradient. As shown in Fig. 1 at x/b=0.4, as the angle of attack

increases, E first decreases, then suddenly increases to a certain

value and finally decreases. If s and r are the minimum and maximum

points, respectively, point s represents the beginning at which

unstable flow becomes obvious, and point rcorresponds to the complete transition of the flow into a turbulent one, i.e., the transition point. Change in E with angle of attack is also shown in

Fig. 1, where each curve has a maximum between s and r, represented

by m. The S, R and M curves were drawn using the s, r and m values

in Fig. 1 and taking x/b and <X as abscissa and ordinate, as shown in

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£(mV)

y 2 * 60

6 a°

-2 6**

Fig. 1. Variation of E and Ermg with angle of attack,

at different x/b. X-output voltage, E;

A-root mean square of the pulse voltage, E

rms'

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ORIGINAL PAGE IS

OF POOR QUALITY

(a ) 1

Fig. 2. The oscillograms of the output voltage of the

surface hot film gage at x/b=0.7.

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040. 5 oTe O.T o.s */&

Fig. 3. The location of the boundary layer transitions

determined in the experiments.

1-turbulent flow; 2-transition; 3-laminar flow.

Fig. 3. From these three curves, the transition region of the airfoil at different angles of attack can be determined. The results

are generally consistent with the rule which states that the transition point moves ahead as the adverse pressure gradient increases.

The data points, x and + in Fig. 3 are from the exterior insertion /404

values in the Appendix of reference C1] , and reference [4] , which are

generally consistent with the results obtained in this study.

5. CONCLUSIONS

(1) Surface hot films are promising elements capable of

locating the boundary layer transitions.

(2) Location of the boundary layer transitions by using the

multilayer surface hot films and measuring the output voltage of

the CTA circuit, the pulse voltage root mean square, E and the

mis

voltage oscillogram is a feasible and practical technique.

ACKNOWLEDGEMENT

The author thanks Professors Yang Zasheng and Zhang Peilin

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for reviewing the original manuscript and giving very helpful

suggestions, and Xu Xinxue, Shen Hongxia, Zheng Faming and Shen

Fengxiang for the experimental work.

REFERENCE

£l] ^ McCrosky, W J.and Durbin, E J, Flow Angle and Shear Stress Measurements Using Heated

Films and Wires, AIAA i Vol 8, No 3(1970), pp 518~523.

f 2] m Meier, H U , Kreplin, H.-P and Faag, L W , Experimental Study of Two- and Three-Dimensional Boundary Layer Separation, DFVLR IB222 81A12U981)

15], Wang Tiecheng, Principles, Structure and Applications of Surface hot films, in A Catalogue of the Symposium Procedings,

Chinese Aeronautical Society, No. 8215001(1982).

,. -I Lin, C.C., Turbulent Flow and Heat Transfer, Princeton University Press(1959), P. 58.

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