28TH INTERNATIONAL CONGRESS OF THE AERONAUTICAL SCIENCES
PERFORMANCE ANALYSIS OF A SCRAMJET INLET-
ISOLATOR USING EXPERIMENTAL & NUMERICAL
METHODS
A. Che Idris, M. R. Saad, H. Zare-Behtash, E. Erdem & K. Kontis
University of Manchester
Azam.cheidris@postgrad.manchester.ac.uk
Abstract though a scramjet aircraft would certainly
change its angle of attack sometimes during its
Experimental and numerical investigations were overall flight trajectory. Thus this paper seeks to
conducted to understand the complex flow contribute to the overall body of knowledge by
features inside a generic two dimensional conducting experimental and numerical
planar scramjet inlet-isolator. The inlet-isolator investigations relating to the subject matter.
was designed to satisfy the ideal condition of Three most common scramjet inlet
Shock-on-Lip (SOL) in Mach 5 inviscid flow. performance indicators are Total Pressure Ratio,
Angle of attack was varied to simulate off- Kinetic Energy Efficiency and Dimensionless
design flight conditions for the inlet-isolator. Entropy Increase [9]. Total Pressure Ratio
Numerical methods were used to estimate shows the fraction of total pressure still
various performance indicators at different test available in the flow after the compression
conditions. The changes in inlet efficiencies process. Typically a hypersonic inlet will
were shown to be heavily influenced on the flow exhibit lower value of Total Pressure Ratio if
quality inside the isolator. compared to a supersonic inlet [9].
Kinetic Energy Efficiency on the other
hand demonstrates the potential flow velocity
1 Introduction
obtainable if the compressed flow in throat
For practical reason, a scramjet inlet-isolator is section were to be isentropically expanded
optimised for operation at one specific flight without combustion. It is referenced to
condition which is usually the cruise condition freestream flow velocity and usually has a very
where the aircraft will spend most of its flying close value to unity [9].
time. However at off-design conditions, many Since severe interactions of shocks and
adverse flow conditions could occur whereby boundary layer are expected in a hypersonic
reducing its performance or even inducing inlet inlet, a good measure of irreversibility is done
un-start. by calculating the Dimensionless Entropy
Bachchan and Hillier [1] tries to classify Increase. A high quality flow would have a low
different inlet off-design conditions into five value of entropy increase [9].
main types depending on the combinations of Using a validated numerical simulation
shocks from forebody and cowl segment. Many to calculate the inlet performance simplifies
studies has been conducted to investigate the inlet parametric studies without the need to
relationship between inlet flowfield and perform many invasive measurement techniques
different off-design variables such as flight and experiments.
Mach number [1], angle of attack (AoA) [2],
yaw angle [3], and freestream air temperature
[4][5][6][7][8].
However, there is a serious lack of 2 Experimental Setup
references discussing the effect of angle of
attack on scramjet inlet performance even 2.1 High Supersonic Tunnel
1
A. CHE IDRIS et. al.
The experiments were carried out in High experiments performed using the HSST [12].
Supersonic Tunnel (HSST) producing Mach 5 For the range of Reynolds number obtainable
flow with Reynolds number of 13.2 × 106 m−1. from the HSST, laminar flow is expected unless
The blow down tunnel has stagnation pressure tripped thus transitional flow option is selected
of 6.50 bar (±0.05 bar) and stagnation in the solver [12].
temperature of 375 K (±5 K). The setup is Air is specified as ideal gas and
similar to the one used in reference [10] (see Sutherland Law is used to calculate the
Fig. 1). viscosity. The boundary condition at Freestream
and Inflow are specified using HSST test
2.2 Scramjet Inlet-isolator Model section’s condition whilst the Outflow
properties are extrapolated from the interior.
The intake model consisted of a double ramp Constant temperature is used at the solid walls.
with see through cowl segment. It was designed Inviscid solution was used as the
to satisfy Shock-on-Lip condition in inviscid initialization value for the turbulence
Mach 5 flow. The overall length, height and simulation. Grid density sensitivity was
width of the inlet-isolator model is 155 mm × investigated by using coarse (34,485), medium
33.6 mm × 36 mm. (52,250) and fine (73,840) grid.
In this investigation, inlet-isolator model
was subjected to three different angle of attacks
of 0 (baseline), 2 and 4 degree respectively. 3 Results and Discussions
10 pressure tappings are installed
alongside the middle line of the model 3.1 Sensitivity Analysis
connected to Kulite Static Pressure Transducer.
The material used for double ramp and From Fig. 3 we can conclude that numerical
cowl frame is aluminium alloy while the cowl results from Medium and Fine mesh are very
window segments are made from Perspex. The similar. Coarse mesh also produces almost
digram is shown in Fig. 2. similar pattern of pressure peaks in the isolator
section (downstream of x = 0.08 m) but the
2.3 Colour Schlieren Setup mesh is not fine enough to resolve flow
separation around compression corner between
The schlieren system used is a Toepler’s z-type first and second ramp. The solution
such as the one used in [11]. The system is demonstrates grid independent thus allowing the
consisted of a 300 W continuous Xenon Arc use of Medium grid for all case studies reported
Lamp with focusing lens and a 2mm wide slit, herein.
two parabolic mirrors, tri-coloured knife edge
and a set of Hoya 49mm lenses for image 3.2 Numerical Accuracy
focusing. The camera used is Canon EOS-450D
12 MP set at continuous shot of 3.5 frames per Normalized static pressure readings along
second. centerline of the scramjet inlet are plotted in
Fig. 7. For all AoA, excellent agreements are
present between numerical and experimental
2.4 Numerical Methods
data. Qualitatively, density gradient contours
Favre Averaged Navier-Stokes (FANS) plot from numerical simulation are observed to
equations are solved by using density based match very closely the flow topology captured
commercial solver Fluent. Second order by colour schlieren (see Fig. 4, Fig. 5 and
spatially accurate scheme are utilized together Fig.6).
with Roe’s Flux-Difference Splitting.
Menter’s Shear-Stress Transport κ-ω are 3.3 Baseline Case
selected as the main turbulence model as it has
been repeatedly validated by previous
2
PERFORMANCE ANALYSIS OF SCRAMJET INLET-ISOLATOR
USING EXPERIMENTAL AND NUMERICAL METHODS
Fig. 4 shows complex and comprehensive view 3.4 Effects of AoA on Inlet-isolator
of flow topology for a Mach 5 scramjet inlet- Flowfield
isolator with zero degree angle of attack. There
are four separation bubbles identified by With increase in AoA, the separation at
presence of separation and reattachment shocks. compression corner decrease in size (see Fig. 4,
Separation at compression corner Fig. 5 and Fig.6). The compression corner’s
produced Goetler vortices that assist in separation shock appeared ‘smeared’ as the
transition to turbulence [13]. Separation and separation become more subtle. Reference [18]
reattachment shocks from the separation bubble has observed that separation at compression
intersect each other to form Edney’s Type VI ramp became large with unit Reynold’s number.
shock interaction. Since both of Separation In our case, lower flow Mach number
Shock and Reattachment Shock are from the experienced by both ramps at higher AoA,
same family (i.e. have the same shock reduce the unit Reynolds number. Cowl tip
direction), a stronger shock and expansion wave separation also becomes smaller with increasing
is produced from the point of interaction with AoA.
the flow divided by a slip line. Thus flow As the inlet increase it AoA the total
downstream of separation is non uniform. turning of the flow before it enters the throat
Separation Bubble 1 originating from also increase thus lowering the entrance Mach
cowl lip is thought as a product of lip bluntness number. Low Mach number combined with
effect [14]. As lip cannot be made perfectly ‘tall’ shoulder separation bubble in 4 deg AoA
sharp due to manufacturing constraint, bow case enabled the formation of Edney’s Type II
shock emerged and gave rise to entropy layer shock-shock interaction (see Fig. 6). It is similar
with strong vorticity [15]. Shock from to Edney’s Type I interaction where two shock
Separation Bubble 2 interacted with this layer of different family intersect each other but of
resulting in a bubble of recirculated flow [14]. It such strength that a Mach stem must exist
is an inviscid phenomenon but could appear as between the two shock to enable the required
boundary layer separation in real life viscous flow turning. Downstream of the Mach stem is a
fluid. Lip bluntness effect is not present in the subsonic pocket bounded by two slip lines
numerical colour schlieren in Fig 4(a). This emanating from the intersections of the Mach
explains the slight discrepancy in pressure stem with the two oblique shocks. Thus the flow
reading between experiments and simulation in is highly non-uniform around the throat area
Fig. 7. with some part has subsonic speed while the
Interaction between Separation Shock 1 other stays supersonic. However as we can still
and 2 is identified as Edney’s Type 1. Since observe fully established oblique shock train
both of the shock is from different family, it can downstream of throat the subsonic flow
cross each other. Thus two transmitted shock eventually became supersonic again after it pass
emerged from the intersection point. Flow the expansion fans around the two separation
downstream of a transmitted shock has different bubbles.
properties except pressure and velocity from the The shock train inside the isolator
flow downstream of another. This will add to section becomes more ‘compact’ as the shocks
flow non-uniformity and could also cause the reflect in shorter distance with increase in AoA.
flow to become unsteady due to presence of
Kelvin-Helmholtz instability at the slip line. 3.5 Effects of AoA on Inlet-isolator
[16]. Performance
The inlet-isolator is prone to unstart due
to presence of Separation Bubble 1 and 2 which Flow properties from the simulation were mass-
increased the contraction ratio. Unstart will averaged to calculate Total Pressure Ratio,
occur if contraction ratio goes beyond Kinetic Energy Efficiency and Dimensionless
Kantrowitz limit [17]. Entropy Increase using equations from [9].
3
A. CHE IDRIS et. al.
From Fig. 8, Fig. 9 and Fig. 10, it is [1] Bachchan N and Hillier R. Hypersonic Inlet Flow
apparent that scramjet inlet-isolator performance Analysis at Off-design Condition. AIAA Paper 2004-
5380, 2004
improves slightly with increase in AoA. [2] Boon S and Hillier R. Mach 6 Hypersonic Inlet Flow
However, if the AoA is such that a Mach stem is Analysis at Incidence. AIAA Paper 2006-3036, 2006
introduced into the flow, the overall [3] Hohn O and Gulhan A. Experimental Investigation
performance then dropped quite dramatically. of the Influence of Yaw Angle on the Inlet
The improvements in performance by Performance at Mach 7. AIAA Paper 2010-938, 2010
[4] MacRae DS and Neaves M. Time Accurate
the inlet-isolator at AoA = 2 compared to at Computations of Unsteady Hypersonic Inlet Flows
AoA = 0 are due to smaller flow separation at with a Dynamic Flow Adaptive Mesh. Airforce
compression corner and at cowl tip region. Flow Office of Scientific Research Final Technical Report,
separation introduces unnecessary separation 1998
and reattachment shocks which increase the [5] Mayer DW and Paynter GC. Prediction of Supersonic
Inlet Unstart Caused by Freestream Disturbances.
total pressure loss and promotes flow non- AIAA Journal, Vol. 33, No. 2, pp 266-275, 1995
uniformity. Smaller separation with weaker [6] Chang J, Bao W, Yu D, Fan Y, and Shen Y. Effects
separation and reattachment shocks will of Wall Cooling on Performance Parameters of
improve the performance. Hypersonic Inlets. Acta Astronautica, Vol. 65, pp
On the other hand, appearance of Mach 467-476, 2008
[7] Lin KC, Tam CJ, Eklund DR, Jackson KR, Jackson
stem (normal shock) in the case of AoA = 4 and TA. Effects of Temperature and Heat Transfer on
lower the performance of the inlet significantly. Shock Train Structures Inside Constant Area
A normal shock is exactly the shock phenomena Isolators. AIAA Paper 2006-817, 2006
most unwanted inside a scramjet inlet. It [8] Krause M, Reinartz B and Ballmann J. Numerical
produces unbearable temperature increase and is Investigation and Simulation of Transition Effects in
Hypersonic Intake Flows. In High Performance
highly inefficient in compressing air at Computing in Science and Engineering, pp 391-406,
hypersonic speed. The presence of a subsonic 2007
pocket downstream of the normal shock [9] Heiser WH and Pratt DT. Hypersonic Airbreathing
increases flow non-uniformity which will Propulsion. 1st edition, American Institute of
adversely affect combustion performance. Aeronautics and Astronautics, 1994
[10] Erdem E, Yang L and Kontis K. Drag Reduction
Studies by Steady Energy Deposition at Mach 5.
AIAA Paper 2011-1027. 2011
4 Conclusions [11] Yang L, Zare-Behtash H, Erdem E, and Kontis K.
Application of AA-PSP to Hypersonic Flows: The
The study demonstrates the ability of Fluent to Double Ramp Model. Sensor and Actuators B:Chem.
2011
closely predict the flow features of a scramjet [12] Erdem E. Active Flow Control Studies at Mach 5:
inlet-isolator at Mach 5; closely matching the Measurement and Computation. PhD Thesis.
experimental observation. This has enabled a University of Manchester, 2011
rapid performance comparison of the inlet at [13] Saric WS. Goetler Vortices. Ann. Rev. Fluid Mech.
various off-design conditions showing the 26, 1994
[14] Boon S and Hillier R. Hypersonic Inlet Flow
general relations between flow topology and the Analysis at Mach 5, 6 and 7. AIAA Paper 2006-12,
calculated performance indicators. 2006
The study also shows how significant is [15] Anderson JD. Hypersonic and High Temperature
the effect of flow separations to the inlet Gas Dynamics. American Institute of Aeronautics
performance. Highly detailed colour schlieren and Astronautics, 2006
[16] Bertin JJ. Hypersonic Aerothermodynamics.
images accompanied by matching numerical American Institute of Aeronautics and Astronautics,
simulations are invaluable to an inlet designer in 1994
determining how and where to implement flow [17] Das S and Prasad JK. Starting Characteristics of a
control for the inlet. Rectangular Hypersonic Air Intake with Cowl
Deflection. The Aeronautical Journal, 114, pp 177-
189, 2010
[18] Matsumura S, Schneider SP and Berry SA.
Streamwise Vortex Instability and Transition on the
References
4
PERFORMANCE ANALYSIS OF SCRAMJET INLET-ISOLATOR
USING EXPERIMENTAL AND NUMERICAL METHODS
Hyper-2000 Scramjet Forebody. Journal of
Spacecraft and Rockets, 42, 2005
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proceedings or as individual off-prints from the
proceedings.
Fig. 1. High Supersonic Wind Tunnel (HSST) setup [10]
Fig. 2. Schematic diagram of scramjet inlet-isolator model
5
A. CHE IDRIS et. al.
Fig. 3. Plot of normalized pressure of scramjet inlet-isolator model using coarse, medium and fine mesh
Separation
bubble 1
Ramp
Shock
Separation
bubble 2
Compression
corner
separation
Separation
Ramp bubble 1
Shock
Separation
bubble 2
Compression
corner
separation
Fig. 4. [Top] Colour schlieren image of baseline case. [Bottom] Numerical colour schlieren of baseline case
6
PERFORMANCE ANALYSIS OF SCRAMJET INLET-ISOLATOR
USING EXPERIMENTAL AND NUMERICAL METHODS
Separation
bubble 1
Ramp
Shock
Separation
bubble 2
Compression
corner
separation
Separation
bubble 1
Ramp
Shock
Separation
bubble 2
Compression
corner
separation
Fig. 5. [Top] Colour schlieren image of AoA = 2deg case. [Bottom] Numerical colour schlieren of AoA = 2deg case
Separation
bubble 1
Ramp
Shock
Separation
bubble 2
Compression
corner
separation
Separation
bubble 1
Ramp
Shock
Separation
bubble 2
Compression
corner
separation
Fig. 6. [Top] Colour schlieren image of AoA = 4deg case. [Bottom] Numerical colour schlieren of AoA = 4deg case
7
A. CHE IDRIS et. al.
Fig. 7. Normalized static pressure along centreline of scramjet inlet-isolator for baseline, AoA = 2deg and AoA = 4 deg
Fig. 8. Effect of AoA on Total Pressure Ratio
Fig. 9. Effect of AoA on Kinetic Energy Efficiency
8
PERFORMANCE ANALYSIS OF SCRAMJET INLET-ISOLATOR
USING EXPERIMENTAL AND NUMERICAL METHODS
Fig. 10. Effect of AoA on Dimensionless Entropy
Increase
