The experimental investigation on a 600 apex-angle is carried out to study feasibility of simultaneous measurement of drag and aerodynamic heating in a shock tunnel. A three-component accelerometer balance system is used to measure the drag whereas platinum thin film gauges are used for heating rate measurements on the surface of the blunt cone model. Since every ground test and flight test data have unique test conditions, so the importance of simultaneous measurement is realized by measuring both drag and heating rates under unique test conditions during gas injection at the nose of a blunt cone model flying at Mach 5.75, in a IISc hypersonic shock tunnel (HST2). In addition, flow visualization experiments are also conducted on the blunt cone model using electric discharge technique

Sahoo, N.

University of Queensland eSpace

16th Australasian Fluid Mechanics Conference 
Crown Plaza, Gold Coast, Australia 
2-7 December 2007 
 
Experiments on a Blunt Cone Model in a Hypersonic Shock Tunnel 
 
Niranjan Sahoo* 
*Assistant Professor, Department of Mechanical Engineering 
Indian Institute of Technology Guwahati, 781 039, INDIA 
Email: shock@iitg.ernet.in  
        
Abstract 
The experimental investigation on a 600 apex-angle is carried out 
to study feasibility of simultaneous measurement of drag and 
aerodynamic heating in a shock tunnel. A three-component 
accelerometer balance system is used to measure the drag 
whereas platinum thin film gauges are used for heating rate 
measurements on the surface of the blunt cone model.  Since 
every ground test and flight test data have unique test conditions, 
so the importance of simultaneous measurement is realized by 
measuring both drag and heating rates under unique test 
conditions during gas injection at the nose of a blunt cone model 
flying at Mach 5.75, in a IISc hypersonic shock tunnel (HST2). 
In addition, flow visualization experiments are also conducted on 
the blunt cone model using electric discharge technique.   
 
Introduction 
 
The aerodynamic drag and heating are the two dominating 
features in the design of hypersonic vehicles. The aerodynamic 
heating increases as the cube of the velocity whereas the 
aerodynamic drag increases as the square of the velocity [1]. The 
heating problems are reduced to some extent by employing the 
large angle body configurations, but then the drag increases.  In 
the context of model geometry, the stagnation point aerodynamic 
heating varies inversely as the square root of nose radius [1]. So 
the larger is the nose radius, lower is the aerodynamic heating 
and vice versa. Many variations of the aerodynamic shapes were 
studied to address the issues of heating and drag. Among the 
most recent advances, is the Jupiter re-entry vehicle (45o semi-
apex angle blunt cone model with bluntness of 0.352) and Mars 
pathfinder vehicle (70o apex-angle sphere-cone model with 
bluntness of 0.5). The ballistic range of tests conducted on Jupiter 
entry vehicle configuration (Galileo probe) highlight the 
importance of aerodynamic drag [2] whereas the Mars entry 
configuration [3] highlights the importance of aerodynamic 
heating tests. The surface heat transfer rates over 1400 blunt cone 
configurations are also routinely measured in impulse types of 
testing facilities [4, 5]. These reports in the literature address the 
importance of aerodynamic heating and drag independently 
because the model size doesn’t allow implementing requisite 
instrumentation for simultaneous measurements. Moreover, 
unique test conditions can not be repeated in short duration 
testing facilities if the drag and heating rates are measured 
independently.  
 Severe aerodynamic heating during re-entry, is another 
concern for designing hypersonic vehicle which in turn decides 
the requisite thermal protection system (TPS). Hence, 
simultaneous measurement of aerodynamic heating and drag on 
blunt cone models with a feasible thermal protection will address 
some of the issues in a hypersonic flow fields. A novel method 
was proposed by Menezes et al. [6] by using a forward facing 
retractable aero-spike for reducing the wave drag for blunt-nosed 
bodies flying at supersonic and hypersonic speeds. It was shown 
that by incorporating suitable geometry of aero-spikes the flow 
field around the blunt cone can be modified. However, the spike 
becomes hot and ablates as a result of large stagnation 
temperature, hence requires frequent replacement or active 
cooling. One of the alternate solutions is to achieve similar effect 
using a forward facing gaseous jet. In the present paper, the 
experimental investigations are carried out on a 600 apex-angle 
blunt cone model to explore the effectiveness of forward facing 
gaseous jets on aerodynamic flow field modification where the 
effect of drag and heating rates are predominant under a given 
test condition. In particular, the model is designed such that both 
drag and surface heat transfer rates can be measured 
simultaneously. In addition, flow visualization study using 
electric-discharge technique is attempted to understand the shock 
structures in the hypersonic flow field. All the experiments are 
carried out at flow Mach number of 5.75 in IISc hypersonic 
shock tunnel HST2. Moreover, the data for surface heating rates 
and drag are presented for unique test conditions which not only 
cost effective but also saves time. 
  
Experimental Facility 
 
The experimental investigations are carried out in the HST2 
hypersonic shock tunnel, shown schematically in Figure 1.  
 
1 2 3 4
5
6
Computer 
based data 
acquisition 
systemU
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C
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Dump 
tank
Vacuum 
pump
Exhaust
78 8
9
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1. Driver section
2. Aluminum diaphragm
3. Driven section
4. Paper diaphragam
5. Conical nozzle
6. Test section
7. Location of pressure transducer
8. Thermal sensors ports for shock speed
9. Driven section vacuum pump
10. Pressurized driver gas cylinder
All dimensions are in 'cm'.
 
Figure 1: Layout of IISc hypersonic shock tunnel, HST2. 
 
It comprises of a shock tube, a conical nozzle, square test section 
and a dump tank along with a high efficiency vacuum pumping 
system. The tunnel is capable of producing a reservoir enthalpy 
up to 4.0 MJ/kg and has an effective test time of about 800 µs. 
The hypersonic flow goes through a 450mm long test section of 
300mmx300mm square cross section before entering the 
1000mm diameter dump tank. The driver and driven sections of 
the shock tube are separated by a metallic diaphragm of 
appropriate thickness and the shock tube and the nozzle are 
separated by a thin paper diaphragm. Typical test conditions 
achieved in HST2 shock tunnel are listed in Table 1. 
 
 
 
503
Table 1: Nominal test conditions in HST2 
 
Driver gas Nitrogen and 
helium  
Primary diaphragm thickness  1mm and  
1.5 mm 
Shock Mach number,  
Ms(± 2.5%) 
2.1 to 3.6 
Stagnation pressure, 
 Po  (± 8%) 
220 to 1030 kPa 
Stagnation enthalpy 
 Ho (± 10%) 
0.70 to 1.6 
MJ/kg 
Free stream Mach number 
 M∞  (±2%) 
5.75 
Free stream static pressure 
 P∞  (± 5%) 
0.18 to 0.85 kPa 
Free stream static temperature, 
T∞ (± 9%) 
90 to 210 K 
Free stream unit Reynolds 
number, Re∞ (± 16%) 
~ 106  (m-1) 
 
 
Measurement of Drag and Aerodynamic Heating Rates 
 
The phenomenon of forward facing gaseous jet as a means of 
flow field modification is explored by measuring drag and 
surface heating rates around a blunt cone model flying at Mach 
5.75 in HST2 shock tunnel at 00 angle of attack.  The 600 apex-
angle blunt cone with bluntness of 0.857 is shown in Figure 2 
whereas the photograph of the model is shown in Figure 3. 
1. Drag accelerometer
2. Rubber bush
3. Stainless steel ring
4. Central sting
5. Macor strip
6. Platinum thin film sensors
All Dimensions are in 'mm'.
1
2
6
5 5
33
42
 
Figure 2: A 600 apex-angle blunt cone model fitted with an 
accelerometer balance system and platinum thin film gauges for 
simultaneous drag and heat transfer measurements  
 
 
 
Figure 3: Photograph of the 600 apex-angle blunt cone model 
along with an accelerometer balance system and platinum thin 
film gauges for simultaneous drag and heat transfer 
measurements.  
The gases under atmospheric conditions are injected into the 
hypersonic flow through a hole of 2 mm diameter at the nose of 
the model. A flexible pipe connected to this hole passes internally 
along the hollow sting that in turn is connected to the tunnel port. 
The aerodynamic drag over the model is measured using the 
three-component accelerometer balance system whereas the 
platinum thin film sensors deposited on the Macor insert 
measures convective heating rates. The basic principle of 
measurement of aerodynamic forces and surface heating rates are 
highlighted by Sahoo et al. [7] and Jagadeesh et al. [5], 
respectively. Because of having a hole of 2 mm diameter at the 
nose of the model, it is almost impossible to mount the heat 
transfer gauges at the stagnation point. However, the heat transfer 
gauges mounted closest to the stagnation point is at the location 
of ( )nRs  as 0.3333 where s  is the arc length on the cone 
surface measured from geometric stagnation point and nR  is nose 
radius. This is again based on the minimum thickness of the 
Macor for which one-dimensional heat conduction equation can 
be assumed. 
The test model (Figure 2 and 3) is used to study the 
phenomena of forward facing gaseous jet during shock tunnel 
testing at 00 angle of attack under two different stagnation 
enthalpy conditions at Mach 5.75. Four different gases (air, 
argon, carbon dioxide and helium) are injected into the 
hypersonic stream through a hole of 2 mm diameter at the nose of 
the model. The operating pressure ratio ensures choked mass 
flow rate at the nozzle exit plane.  
A typical temperature-time history and the corresponding 
numerically integrated heat transfer signals with and without gas 
injection are shown in Figure 4 and 5, respectively.  The force-
time history inferred from the drag accelerometer with and 
without air injection is shown in Fig. 6.  
 
Figure 4: Typical temperature-time history from platinum thin 
film sensors ‘with and without’ gas injection. 
 
 Maximum reduction in heating rate of about 40 % is 
observed at the gauge location of ( )nRs = 0.7333 for stagnation 
enthalpy of 1.16 MJ/kg with carbon dioxide injection and the 
corresponding values change to 28 % with helium injection. 
About 12-25 % increase in the aerodynamic drag coefficient is 
observed at all enthalpy conditions. The reason for the drag 
enhancement is because the operating pressure is not adequate 
enough to produce a gas jet into the oncoming hypersonic flow 
and thereby forms a thin film around the test model. It leads to 
‘heat buffer’ between the hot air stream and the body which 
reduces the heat transfer rates substantially. Since, the operating 
pressure ratio is not sufficient to push the body shock away from 
the blunt body surface, and hence leads to increase in the gas 
density due to gas injection in the stagnation region which in turn 
increases aerodynamic drag. The detail investigation on 
504
effectiveness of forward facing gaseous jet on film cooling can 
be found in Sahoo et al. [8]. 
 
Figure 5: Numerically integrated heat transfer signal for platinum 
thin film sensors ‘with and without’ gas injection. 
 
 
 
Figure 6: Typical force-time history inferred from drag 
accelerometer at 00 angle of attack.  
 
 
Flow Visualization Study for Blunt Cone Model 
 
In order to highlight the factual evidence of shock structure with 
forward facing gaseous jet into the hypersonic flow field, flow 
visualization study around the blunt cone model is attempted. 
This would essentially complement the results of simultaneous 
measurement of drag and surface heating rates. The electric 
discharge technique [9] is used in the present study to visualize 
shock structure on the blunt cone configuration. This technique is 
based on the principle that the intensity of spontaneous radiation 
emitted by the ion recombination in an electric discharge depends 
on the gas density and temperature along the discharge path. So 
when an electric discharge is generated across a shock wave, the 
light emitted from the shock wave region will be very weak 
compared to that of free stream and shock layer. Based on the 
density and temperature difference of the shock layer and free 
stream, the shock shapes appear in the intensity field and can be 
captured by taking the photograph of the discharge column.  
The schematic diagram of the electrode arrangement for the 
600 apex-angle blunt cone with 150 flared model, used for flow 
visualization studies is shown in Figure 7. A 1 mm diameter 
stainless steel sewing needle with a sharp tip suspended from the 
roof of the test section is used as a point electrode (+ve) while a 
0.5 mm thick copper strip, embedded into the model such that 
one of its edges is flush with the front surface of the blunt cone, 
acts as a line electrode (ground). The electrodes are kept about 50 
mm apart and operating voltage and current across the electrodes 
are fixed at 1.8 kV and 1Amp, respectively. A Bakelite hylem 
blunt cone model is used in the experiments to ensure insulation 
and also to avoid multiple reflections from the model surface. 
The application of high DC voltage across the electrodes is 
controlled by the switching transistor, which in turn is controlled 
by the triggering pulse from the pulse delay unit [10]. In most of 
the experiments the duration of the electric discharge between the 
electrodes is maintained at ~ 2 µs. In order to avoid blockage of 
the flow in the tunnel test section and to maintain sufficient gap 
between the model edge and the point electrode attached to the 
test section wall, the size of the flow visualization test model is 
scaled down to 42 % of the one used for force and heat transfer 
measurements while maintaining the geometrical similarity. A 
SLR camera (Cannon F1.4) operated in the bulb exposure mode 
using ASA 1600 speed film, which is further boosted to ASA 
6400 by appropriate developing technique, is used to capture the 
photograph.  A typical experimental result of the shock shape 
visualization carried out for the enthalpy conditions of 1.16 
MJ/kg is shown in Figure 8 without gas injection.  
 
Point electrode
Line electrode
Free stream at 
Mach 5.75
Tunnel test section wall
Tunnel test section wall
All dimensions are in 'mm'.
 
Figure 7:  Schematic diagram of blunt cone model with electrode 
arrangement for flow visualization experiment. 
 
 
1
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3
4
5
H
yp
er
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flo
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  M
in
f =
 5
.7
5
1. Blunt cone model
2. Line electrode
3. Bow shock
4. Point electrode
5. Tunnel test section wall
 Shock angle = 39 deg.
 
Figure 8: Shock structure around the blunt cone model.  
 
In this picture, the bright lines in the conical portion of the 
model indicate the shock position. However, the bright lines are 
not clear in the nose region because the discharge from the point 
electrode in the test section didn’t strike the line electrode 
embedded on the nose portion of the model surface. The portion 
of the body in which the strength of the discharge column is 
505
more, appears as bright line and discharge column gets diffused 
near the spherical portion of the model.   
The motivation of flow visualization study undertaken in the 
current investigation is to observe whether or not the hypersonic 
flow field gets modified due to gas injection. Unfortunately, the 
shock shapes could not be captured with gas injection due to 
several technical difficulties encountered during experiment. In 
fact, the whole process involves the synchronization of electrical 
discharge with the establishment of hypersonic flow in the test 
section with gas injection, controlling the delay and discharge 
times in the pulse delay unit and most importantly photographic 
recording during electric discharge.  So no other details are being 
reported here because this method needs further improvements in 
this area in order to capture shock shapes with gas injection. 
However, the shock shapes around the blunt cone model without 
gas injection confirms the hypersonic flow quality in the tunnel. 
 
Conclusions 
 
Simultaneous measurement of aerodynamic heating and drag is 
carried out for a 600 apex-angle blunt cone model to study the 
effect of hypersonic flow field modification due to a forward 
facing gaseous jet at the nose of the model. Four different gases 
(air, carbon dioxide, argon and helium) are injected into the 
hypersonic flow field at different enthalpy levels. Maximum 
reduction of surface heating rates of 40% is observed in the 
vicinity of stagnation point with carbon dioxide injection whereas 
the aerodynamic drag is enhanced by 12-25%. Flow visualization 
study is attempted to reveal the shock structure in the hypersonic 
flow field. Although shock shapes around the test model without 
gas injection gives a clear picture, but due to experimental 
difficulty, similar experiments could not be completed with gas 
injection.  
 
 
Acknowledgments 
 
First of all, the author would like express heartfelt gratitude to his 
research supervisor Prof. K.P.J. Reddy for carrying out 
experiments in IISc hypersonic shock tunnel HST2. At present, 
the author’s research activities at Australia is funded by 
Department of Science and Technology (DST), Govt of India, 
under BOYSCAST fellowship programme. So, the author deeply 
acknowledges DST, Govt. of India for providing the necessary 
financial support during the period of stay at Australia as well as 
for attending the 16AFMC conference. The author is also 
thankful to Director, Indian Institute of Technology Guwahati, 
for providing academic leave during BOYSCAST fellowship 
period.  
 
 
 
 
 
References 
 
[1] Anderson J.D. Jr Hypersonic and High Temperature Gas 
Dynamics, McGraw-Hill series in Aeronautical and 
Aerospace Engineering, New York, 1989 
 
[2] Peter F.I. and Donn B.K., High-Speed Aerodynamics of 
Several Blunt-Cone Configurations, Journal of Spacecraft 
and Rockets 24 (2), 1987, 127-132 
 
[3] Brian R.H. and John N.P., High enthalpy 
Aerothermodynamics of a Mars Entry vehicle Part 1: 
Experimental Results, Journal of Spacecraft and Rockets 34 
(4), 1997, 449-456 
 
[4] Stewart D.A. and Chen Y.K., Hypersonic Convective Heat -
transfer over 1400 Blunt Cones in Different Gases, Journal 
of Spacecrafts and Rockets 31 (5), 1994, 735-743 
 
[5] Jagadeesh G, Reddy N.M., Nagashetty K. and Reddy K.P.J., 
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[6] Menezes V., Saravanan S., Jagadeesh G. and Reddy K.P.J., 
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[7] Sahoo N., Mahapatra D.R., Jagadeesh G., Gopalkrishnan S. 
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Bodies in a Hypersonic Shock Tunnel, Measurement Science 
and Technology 14, 2003, 260-272 
 
[8] Sahoo N., Vinayak K., Saravanan S., Jagadeesh G. and 
Reddy K.P.J., Film Cooling Effectiveness on a Large Angle 
Blunt Cone flying at Hypersonic Speed, Physics of Fluids, 
17(3), 2005, 1-11.  
 
[9] Nishio M., New Method for Visualizing Three-Dimensional 
Shock Shapes around Hypersonic Vehicles using an 
Electrical Discharge, AIAA Journal 28 (12), 1990, 2085-
2091 
 
[10] Jagadeesh G., Srinivasa Rao B.R., Nagashetty K., Reddy 
N.M. and Reddy K.P.J. Electrical Discharge Technique for 
Three-Dimensional Flow Field Visualization in Hypersonic 
Shock Tunnel, Journal of Flow Visualization and Image 
Processing 4, 1997, 51-57 
 
 
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Experiments on a Blunt Cone Model in a Hypersonic Shock Tunnel

Niranjan Sahoo

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Experiments on a Blunt Cone Model in a Hypersonic Shock Tunnel

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