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adhan
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a Vol.32,Parts1&2, February–April 2007, pp. 123–131. © Printed in India
Experimental investigation of drag reduction by forward
facing high speed gas jet for a large angle blunt cone at
Mach 8
BALLA VENUKUMAR andKPJREDDY
Department of Aerospace Engineering, Indian Institute of Science,
Bangalore 560 012
e-mail: laser@aero.iisc.ernet.in
Abstract. Substantial aerodynamic drag, while flying at hypersonic Mach number,
due to the presence of strong standing shock wave ahead of a large-angle blunt-
cone configuration, is a matter of great design concern. Preliminary experimental
results for the drag reduction by a forward-facing supersonic air jet for a 60
◦
apex-
angle blunt cone at a flow Mach number of 8 are presented in this paper. The
measurements are carried out using an accelerometer-based balance system in the
hypersonic shock tunnel HST2 of the Indian Institute of Science, Bangalore. About
29% reduction in the drag coefficient has been observed with the injection of a
supersonic gas jet.
Keywords. Aerodynamic drag; hypersonic flow; shock tunnel; air jet; force
balance; shock interaction.
1. Introduction
It is well known that in order to address the severe heating problems encountered during the
atmospheric flight of hypersonic vehicles, large-angle blunt-cone configurations are preferred
at the cost of enhanced aerodynamic drag, due to the presence of the nearly normal standing
shock wave in front of the vehicle. The enhanced wave-drag increases the demand for propel-
lant during the ascent phase (Peter & Donn 1987). Thus, the management of the aerodynamic
drag encountered by the blunt body configurations of hypersonic vehicles is of significant
interest in order to reduce the cost of space travel or enhance the range of supersonic mis-
siles. Many techniques have been proposed to reduce the wave drag essentially based on the
aerodynamic drag management by controlling the flowfields around the high speed bodies.
Some of these techniques include the use of aerospikes with different aerodiscs (Menezes
et al 2003), gas injection at the stagnation region (Sahoo et al 2005), multi-step base design
(Menezes et al 2005) and energy deposition into the flow (Satheesh & Jagadeesh 2005). We
have been investigating these techniques experimentally in recent years in the hypersonic
shock tunnel HST2 of the Indian Institute of Science, Bangalore, numerically using the CFX
TASCFlow code. A review of these results was presented at the 11th Asian Congress of Fluid
Mechanics (Reddy et al 2006).
123
124 Balla Venukumar andKPJReddy
Drag reduction by a counterflowing supersonic jet emanating from the stagnation point
of a blunt body flying at hypersonic Mach number is one more powerful technique for the
effective reduction of drag as well as for aerodynamic heat transfer to the body (Josyula et al
2001; Hayashi & Asoy 2003; Balla Venukumar et al 2006). Drag reduction by counterflowing
supersonic jet for a 60
◦
apex angle blunt cone is investigated in the HST2 shock tunnel at a
flow Mach number of 8. Some of the results of this study are reported briefly in this paper.
More detailed results as well as the basic physics of the drag reduction by the counterflow
supersonic jet are presented in a recent publication (Balla Venukumar et al 2006).
2. Experimental facility
The experimental investigations are carried out in the hypersonic shock tunnel HST2 shown
schematically in figure 1. The working of the shock tunnel and the procedure for estimating
its performance are presented in detail in an earlier publication (Reddy et al 1996). The
HST2 shock tunnel consists of a shock tube (50-mm internal diameter) divided into driver
and driven sections separated by a metallic diaphragm. The shock tube is connected to the
wind-tunnel portion consisting of a hypersonic nozzle, test section and dump tank through
a paper diaphragm at the entrance of the nozzle. The hypersonic flow is produced by the
convergent–divergent conical nozzle of 10
◦
divergence angle and 300 mm exit diameter. The
nozzle produces Mach 5·75 flow in the test section in straight-through mode of operation and
hypersonic flow of Mach numbers 8, 10 and 12 are obtained in the reflected mode of operation
by attaching an appropriate throat at the entrance of the nozzle. The experiments described in
this paper are carried out in the reflected mode of operation at Mach 8. The hypersonic flow
goes through a 450 mm long test section of 300× 300 mm square cross section before entering
Figure 1. Schematic diagram of the HST2 hypersonic shock tunnel equipped with the gas injection
system for generating the supersonic gas jet from the stagnation point of the model in the test section.
Drag reduction by gas jet in hypersonic flow 125
Figure 2. (a) Schematic diagram and (b) photograph of a 60
◦
apex-angle blunt cone model fitted
with platinum thin film gauges and accelerometer balance system.
the 1000 mm diameter dump tank. The wind tunnel portion of the shock tunnel is evacuated
to 10
−6
mbar pressure by using a rotary pump–diffusion pump combination. The hypersonic
flow in the test section is calibrated using the pitot rake traversing along the axis parallel to the
flow. The details of the calibration procedure are presented in a previous publication (Reddy
et al 1996).
126 Balla Venukumar andKPJReddy
Figure 3. (Caption on facing page.)
Drag reduction by gas jet in hypersonic flow 127
Figure 3. (a) Photograph of the solenoid valve, (b) circuit diagram for the power supply to the
solenoid valve, and (c) detailed drawing of the injector system.
3. Test model
In the present investigation, a 60
◦
apex-angle blunt cone model with 70 mm base diameter and
bluntness ratio of 0·857 is used. The model is equipped with thin film gauges and internally
mounted accelerometer-based balance system to measure both the convective surface heat
transfer rates over the model surface and aerodynamic drag on the model simultaneously.
The schematic diagram and the photograph of the test model are shown in figures 2a and b
respectively.
Injecting the supersonic jet during the steady hypersonic flow in the test section is the
most challenging task in the present investigations as the total test time in the tunnel is less
than one millisecond. For this purpose we have built a gas-injection system using a solenoid
valve shown in figure 3. The injector system consists of a solenoid valve shown in figure 3a
which is operated by a 24 VDC supplied by the electronic circuit, shown in figure 3b, after
receiving an external trigger signal. The solenoid valve supplies the gas to the orifice at the
stagnation point of the model from the reservoir connected to the gas cylinder as shown in
figure 3c. The operational characteristics of the gas injector system are determined using a test
chamber shown in figure 4a. The test chamber was evacuated to the pressure levels identical
to the levels in the test section and dump tank assembly, and a pressure probe was mounted
opposite to the supersonic jet coming out of the injector system. Typical driver-voltage signal
and pressure signal from the probe are shown in figure 4b. From these signals, it is seen that
the solenoid opens completely in about 15 ms after receiving the driver signal. Thus, in order
to synchronize the counterflowing supersonic jet from the stagnation point of the test model
with the arrival of the hypersonic flow in the test section, it is essential to actuate the solenoid
valve about 15 ms before the establishment of hypersonic flow in the test section.
It is found that the time interval between the bursting of the metallic diaphragm and the
arrival of the test flow in the test section is about 15 ms. Hence, the electronic signal output
from a pressure gauge mounted in the driver section of the shock tube is used to trigger the
