Post on 22-Jan-2022
Vol-4 Issue-4 2018 IJARIIE-ISSN(O)-2395-4396
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CFD ANALYSIS OF HYPERSONIC NOZZLE
THROAT ANALYSIS
Gaurav Kumar1, Sachin Baraskar
2
1Research Scholar, Department of Mechanical Engineering, SOE, SSSUTMS, M.P., INDIA
2Assistant Professor, Department of Mechanical Engineering, SOE, SSSUTMS, M.P., INDIA
ABSTRACT
This paper presents conjugate heat transfer analysis for Mach 12 nozzle of hypersonic wind tunnel. For the
analysis, ANSYS Fluent has been used for both flow and conduction analysis in coupled manner considering actual
material properties. First, steady state simulation has been performed to obtain the settled flow with wall
temperature of 300K. After achieving steady state solution, transient simulation has been performed get
convergence. Flow simulation had done by cooling the throat regime by cool water (25kg/s) which is very normal
flow rate in nozzle flow.
Keyword: - ANSYS Fluent1, Mach number2, Temperature Distribution3, and Nozzle Throat4.
1. INTRODUCTION
A nozzle (from nose, meaning 'small spout') is a tube of varying cross-sectional area (usually
axisymmetric) aiming at increasing the speed of an outflow, and controlling its direction and shape. It is 7m
long nozzle; divided into 8 segments from ease of fabrication and inspection point of view.
Convergent portion (Subsonic portion) is 1772mm long and rest is divergent portion (supersonic portion). Throat is
made of Beryllium Copper alloy with a thickness of 6mm. In order to cool the nozzle, a gap of 5mm is maintained
by using a split throat made of SS304L and through this gap, water is circulated. Adjacent regions of split throat are
made of 15-5PH. Outer shell in throat region is SS304L. The outer shell of subsonic section-2 is SA516 gr 70. To
withstand the thermal profile of the flow, Inconel 617 and Cera blanket are used in the subsonic sections. Nozzle
flow inlet diameter is 270 mm. Material for the divergent sections 2 to 6 has been selected as SS304L with
maximum section length of 1150mm and minimum of 750mm. Physical nozzle exit diameter is 1000 mm.
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DETAIL A
Fig. 1 Geometrical details of the nozzle
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2. RESULTS AND DISCUSSIONS
2.1 Steady State Simulation
The results from steady state simulations are shown in Fig. 2 below. The Mach number, static pressure, and static
temperature distribution along the length of the nozzle axis are plotted in Fig. 2a and Fig. 2b respectively.
-1 0 1 2 3 4 5 6 70
2
4
6
8
10
12
Along the length of nozzle (m)
Mac
h N
umbe
r
Mach no
Fig- 2 a : Mach number distribution
-1 0 1 2 3 4 5 6 70
1
2
3
4
5
6
7
8
9
10x 10
6
Along the length of nozzle (m)
Sta
tic P
ress
ure
(Pa)
Static Pressure
Fig-2b: Static Pressure & temperature distribution
-1 0 1 2 3 4 5 6 70
200
400
600
800
1000
1200
1400
Along the length of nozzle (m)
Sta
tic T
empe
ratu
re (K
)
Static Temperature
Fig- 2c: Static Pressure & temperature distribution
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Fig-2 d: Mach contour
Case-1: Throat thickness is 6mm
Fig. 8a Static Temperature distribution along the Length of the nozzle
Fig-8a: Static temperature distribution along the axis
Fig. 8b Heat Flux distribution along the Length of the nozzle
-1 0 1 2 3 4 5 6 7200
300
400
500
600
700
800
900
1000
1100
1200
Position (m)
Satic
Tem
pera
ture
(K)
Satic Temperature Distribution along the axis
t = 0 sec
t = 1 sec
t = 10 sec
t = 20 sec
t = 30 sec
t = 40 sec
-0.12 -0.1 -0.08 -0.06 -0.04 -0.02 0 0.02 0.04 0.06 0.08200
250
300
350
400
450
500
550
600
-1 0 1 2 3 4 5 6 7-16
-14
-12
-10
-8
-6
-4
-2
0
2
Position (m)
Tota
l Sur
face
Hea
t Flux
(MW
/m2 )
Heat Flux Distribution
t = 0 sec
t = 1 sec
t = 10 sec
t = 20 sec
t = 30 sec
t = 40 sec
-0.1 -0.05 0 0.05 0.1-16
-14
-12
-10
-8
-6
-4
-2
0
2
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0 1 2 3 4 5 6
x 10-3
350
400
450
500
550
600
Position (m)
Satic T
em
pe
ratu
re (
K)
Satic Temperature Distribution along the axis
t = 1 sec
t = 10 sec
t = 20 sec
t = 30 sec
t = 40 sec
Fig-3 c : Static temperature distribution along the thickness of the nozzle
The temperature in the subsonic region of the nozzle reaches a maximum temperature of about 1120 K. At the
interface of the subsonic section and throat, the temperature falls down to 410 K. Along the axis, the throat
temperature increases, and there is drop in temperature at two locations of throat and this is due to axial gap
provided for thermal expansion. In the throat section of the nozzle, the temperature picks up with duration and
reaches a maximum of 560 K and then starts falling. As it approaches the throat end region, the temperature picks up
and drops again. This increase in temperature is due to change of material from Beryllium copper to 15-5 PH.
Further the temperature falls down to minimum of 420 K at the exit of the nozzle.
Fig.4a: Wall static temperature distribution at different time instants
Fig. 4 b. Wall static temperature distribution at different time instants
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Fig. 4c: Wall static temperature distribution at different time instants
In this boundary condition of the nozzle, the coolant flow is maintained at 25 Kg/s. The wall static temperature
distribution and wall heat flux distribution along the nozzle profile is plotted in Fig. 3a and Fig. 3b respectively.
It can be inferred from Fig. 3a, the steady state is reached at 20 sec duration, after that the rise in temperature is
negligible. The maximum temperature experienced in the throat region is 560 K and in the divergent sections, it is
about 360 K. At the exit of the nozzle, the temperature is 420 K. The maximum heat flux is 10.2 MW/m2.
The static temperature distribution along the thickness of the nozzle is plotted in Fig. 3c. At the end of blow down of
40 sec, the temperature on the inner wall is 560 K and on the outer wall is 370 K.
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Case-2: Throat thickness is 5mm
Fig- 5a Static Temperature distribution along the Length of the nozzle
Fig-5b Heat Flux distribution along the Length of the nozzle
-1 0 1 2 3 4 5 6 7200
300
400
500
600
700
800
900
1000
1100
1200
Position (m)
Sat
ic T
empe
ratu
re (
K)
Satic Temperature Distribution along the axis
t = 0 sec
t = 1 sec
t = 10 sec
t = 20 sec
t = 30 sec
t = 40 sec
-0.12 -0.1 -0.08 -0.06 -0.04 -0.02 0 0.02 0.04 0.06 0.08200
250
300
350
400
450
500
550
-1 0 1 2 3 4 5 6 7-20
-18
-16
-14
-12
-10
-8
-6
-4
-2
0
Position (m)
Tota
l Sur
face
Hea
t Flu
x (M
W/m
2 )
Heat Flux Distribution
t = 0 sec
t = 1 sec
t = 10 sec
t = 20 sec
t = 30 sec
t = 40 sec
-0.12 -0.1 -0.08 -0.06 -0.04 -0.02 0 0.02 0.04 0.06 0.08-20
-18
-16
-14
-12
-10
-8
-6
-4
-2
0
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0 1 2 3 4 5 6
x 10-3
340
360
380
400
420
440
460
480
500
520
540
Position (m)
Satic T
em
pe
ratu
re (
K)
Satic Temperature Distribution along the axis
t = 1 sec
t = 10 sec
t = 20 sec
t = 30 sec
t = 40 sec
Fig-5c: Static temperature distribution along the thickness of the nozzle
The temperature in the subsonic region of the nozzle reaches a maximum temperature of about 1120 K. At the
interface of the subsonic section and throat, the temperature falls down to 410 K. Along the axis, the throat
temperature increases, and there is drop in temperature at two locations of throat and this is due to axial gap
provided for thermal expansion. In the throat section of the nozzle, the temperature picks up with duration and
reaches a maximum of 530 K and then starts falling. As it approaches the throat end region, the temperature picks up
and drops again. This increase in temperature is due to change of material from Berylium copper to 15-5 PH. Further
the temperature falls down to minimum of 400 K at the exit of the nozzle.
Fig. 6. Wall static temperature distribution at different time instants
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Fig. 7. Wall static temperature distribution at different time instants
In this boundary condition of the nozzle, the coolant flow is maintained at 25 Kg/s. The wall static temperature
distribution and wall heat flux distribution along the nozzle profile is plotted in Fig. 5a and Fig. 5b respectively.
It can be inferred from Fig. 5a, the steady state is reached at 20 sec duration, after that the rise in temperature is
negligible. The maximum temperature experienced in the throat region is 530 K and in the divergent sections, it is
about 370 K. At the exit of the nozzle, the temperature is 420 K. The maximum heat flux is 10.8 MW/m2.
The static temperature distribution along the thickness of the nozzle is plotted in Fig. 5c. At the end of blow down of
40 sec, the temperature on the inner wall is 530 K and on the outer wall is 360 K.
CONCLUSION:
In presented work the CHT simulations have been carried for Mach 12 nozzle at operating stagnation pressure of
100 bar and stagnation temperature of 1377 K using ANSYS Fluent. The analysis has been carried out for two-inlet
case with two different throat 6mm & 5mm thicknesses, Maximum material temperature of 550K has been observed
in subsonic portion and throat regime. It has also indicated that maximum temperature of wall is about 500K.
REFERENCES
[1]. Back, L. H., Massier, P. F. and Gier, H. L., "Comparison of Measured and Predicted Flows Through
Conical Supersonic Nozzles with Emphasis on the Transonic Region," American Institute of Aeronautics and
Astronautics Journal, August, 1965.
[2]. Graham, R. W. and Deissler, R. G., "Prediction of Flow-Acceleration Effects on Turbulent Heat Transfer,"
Transactions of the American Society of Mechanical Engineers, Journal of Heat Transfer, Vol. 89, Series C,
No. 4, pp. 371-372, November, 1967.
[3]. Bartz, D. R., "Turbulent Boundary-Layer Heat Transfer from Rapidly Accelerating Flow of ELocket
Combustion Gases and of Heated Air," in Advances in Heat Transfer, Ed. by Irvine, T. F., Jr. and Hartnett, J.
P., Vol. 2, Academic Press, 1965.
[4]. Back, LH., Massier, P. F. and Cuffel, R„ F., "Some Observations on Reduction of Turbulent Boundary-
vLayer Heat Transfer in Nozzles," Jet Propulsion Laboratory, California Institute of Technology, Pasadena,
California, National Aviation and Space Administration. Contract No. NAS 7-100, 1965.
[5]. Back, L. H., Cuffel, R. F. and Massier, P. F., "Influence of Contraction Section Shape on Supersonic
Nozzle Flow and Performance." Jet Propulsion Laboratory California Institute] of Technology, Pasadena,
California, NASA Contract No. NAS 7-100, 1971.
[6]. Back, L. H., Massier, P. F. and Gier, H. L. "Convective Heat Transfer in a Convergent-divergent Nozzle,"
International Journal of Heat and Mass Transfer, Vol. 7, pp. 549-568, 1964.
[7]. Fortini, A. and Ehlers, R. C, "Comparison of Experimental to Predicted Heat Transfer in a Bell-shaped
Nozzle with Upstream Flow Disturbances," NASA TN D-1743, August 1963.
[8]. Stanton, T. E., "The Variation of Velocity in the Neighborhood of the Throat of a Constriction in a Wind
Channel," British Aeronautical Research Council Reports and Memoranda No. 1388, May 1930.
Vol-4 Issue-4 2018 IJARIIE-ISSN(O)-2395-4396
8815 www.ijariie.com 35
Shelton, S. V., "A Study of Two-Dimensional Nozzle Flow," Unpublished report, 1971.
[9]. Technique," Ph.D. Thesis, Renesselar Polytechnic Institute, Troy, New York, June 1970.
[10]. Shapiro, A. H., the Dynamics, and Thermodynamics of Compressible Fluid Flow, Vol. I, the Ronald
Press Company, New York, N. Y., 1953.
[11]. Meyer, Th., "Uber zweidimensionals Bewegungsyorgange in einen Gas, das mit
Ueberschallgeschwindigkeit stromt," V.D.I. Forschungshef t -t Vol. 62, 1968.
[12]. Lighthill, M. J., "The Hodograph Transformation in Transonic Flows," Royal Society of London,
Proceedings, Series A, Vol. 191, pp. 323-351, November 1947.
[13]. Taylor, G. I., "The Flow of Air at High Speeds Past Curved Surfaces," Aeronautical Research Council
Reports and Memoranda No. 1381, 1930.
[14]. Hooker, S. G., Aeronautical Research Council Reports and Memoranda No. 132, 1930.