COMPUTATIONAL INVESTIGATION by iaemedu

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									       INTERNATIONAL Engineering and Technology (IJMET), ISSN
International Journal of Mechanical JOURNAL OF MECHANICAL 0976 –
6340(Print), ISSN 0976 – 6359(Online) Volume 3, Issue 1, January- April (2012), © IAEME
         ENGINEERING AND TECHNOLOGY (IJMET)
ISSN 0976 – 6340 (Print)
ISSN 0976 – 6359 (Online)
Volume 3, Issue 1, January- April (2012), pp. 161-178
                                                                       IJMET
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COMPUTATIONAL INVESTIGATION OF THE EFFECT OF NOSE
    BLUNTNESS ON THE EXTERNAL AND INTERNAL
      COMPRESSION IN A MIXED INLINE INTAKE
          Vasudeo B.Jaware*, Amarjit Singh**and Satishchandra V. Joshi***
              *Rajarshi Shahu College of Engineering, Pune (M.S), India
             **Terminal Ballistic Research Laboratory, Chandigarh, India
               *** D.Y.Patil College of Engineering, Pune (M.S), India



ABSTRACT

A computational study has been carried out to find out the flowfield inside a two
dimensional hypersonic intake. The study has been carried out for a free stream Mach
number of 6.5, a unit Reynolds number 1.6 x 106 per meter with an intake contraction
ratio of 2.88. The purpose of this analysis was to investigate the effect of nose bluntness
on the external and internal compression in a mixed compression inline intake. An intake
models with different nose bluntness has been analyzed using 2-D numerical simulations
based on a commercially available CFD code. The CFD code and the turbulence model
used is validated by comparing the experimental results available in literature with
computational results. Computational results show that the nose bluntness has significant
influence on pressure distribution and skin friction coefficient distribution on intake
components. Nose bluntness affects the shock wave / boundary layer interactions on
intake upper surface significantly. Numerical results also indicate that the nose bluntness
has a significant influence on intake performance parameters.

Keywords: Inline intake, Intake fence, Scramjet, Nose bluntness, Hypersonic flow
Nomenclature

CFD    = computational fluid dynamics
D      = dimensional
F      = distance from vehicle nose to intake entry
FES    = first expansion shoulder

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H     = intake height
M     = Mach number
MFR = mass flow rate
OST = oblique shock theory
R     = Nose radius
R     = reattachment location
S     = separation location
sh    = sonic line height
Sl    = separation length
Subscript
0     = free stream (Station 0)
1     = intake entry (station1)
2     = intake exit (station2)
∞     = free stream
t      = total condition

1. INTRODUCTION
Intake of a hypersonic vehicle has a major role on the overall performance of its air
breathing propulsion system. The main function of an intake is to provide the
homogeneous high pressure flow to the engine with minimum aerodynamic losses.
External compression is performed through forebody and ramp shocks outside the intake
and internal compression through a series of oblique shocks inside the intake. Based on
the plane in which internal compression takes place, two types of intake designs are
possible - one where the internal compression occurs in the same plane as the forebody
compression, and the other where internal compression takes place in a plane
perpendicular to the forebody compression. The former type of intake is called inline
compression intake and the later is termed as sidewall compression intake. The external
compression of the flow before entering an intake essentially takes place in vertical plane
through the forebody bow shock and one or more ramp shocks.
        A sidewall compression intake consists of vertical wedge-shaped surfaces, and
internal compression occurs in horizontal planes. This reduces the total in-plane turning
of the flow required to obtain the desired compression. The thick boundary layer over the
intake upper wall is therefore, less likely to separate because of shock–boundary layer
interaction. However the thick boundary layer on the intake upper wall may anyway
separate due to glancing shock interaction. On the other hand, in an inline compression
intake, internal compression is affected in the vertical plane by horizontal compression
surfaces, namely the intake lower and upper surfaces. The boundary layer on the intake
upper wall at the intake-entrance can be quite thick with respect to the intake height. The
probability of large-scale separation regions at the entrance of the intake due to shock-
boundary layer interactions will therefore greatly increase due to additional turning of the
flow in the vertical plane.
      Most of the open literature available is about sidewall compression intake.
Experimental and numerical studies [1-2-3] show the effect of intake geometric
parameters on the performance and detailed flow structure through sidewall-compression


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intake. Besides various other requirements, intake should start at low Mach number,
operate over a wide Mach number range, and should work efficiently at designed Mach
number [4]. However, fixed geometry intakes can only be used over a relatively narrow
range of Mach numbers while variable geometry intakes can be used over wide range of
Mach numbers with a reasonably good pressure recovery [5]. Compression process in an
inline intake can be approximated to be two-dimensional in comparison to the three-
dimensional in a sidewall compression intake. Three dimensional fixed geometry
sidewall compression intakes have better performance than the two dimensional inline
intakes over a relatively wide range of Mach numbers. However, it has relatively lower
mass flow rate and pressure recovery at design Mach number and more complex
flowfield [6]. Sanator et al [7] conducted experimental investigation to study the effects
of leading edge bluntness on boundary layer separation             using surface pressure
measurements. Their tests investigated that the introduction of leading edge bluntness to
2D-inlet type flows at M∞ = 10.55 and Re∞/cm = 6.1 x 104 promoted separation. This
promotion of separation was attributed to the reduced boundary layer edge Mach number.
Townsend [8] conducted pressure measurements upstream of the hingeline on a blunted
flat plate / flap configuration (d = 5.0 mm) at M∞ = 10.0 and Re∞/cm = 5.0 x 104 ,
showed an appreciable reduction in the extent of the separated flow region when leading
edge bluntness was introduced. The state of the boundary layer was diagnosed as being
entirely laminar. Danial Arnel et al [9] carried out review on laminar and turbulent
boundary layer interaction with shock waves. There are no basic differences between
Laminar and turbulent interactions as far as the overall flow topology is considered. Of
course, turbulent interactions vary from laminar interactions in terms of scales, intensity
of pressure rise and thermal effects. A shock wave turbulent boundary layer interaction
raises difficult problems which are still largely unsolved. Considering above pros and
cons and lack of published work, a study has been carried out to investigate the flow
through an inline intake with different nose bluntnesses. The nose bluntness has
significant influence on pressure distribution and skin friction coefficient distribution on
different parts of the intake. Nose bluntness affects the shock wave / boundary layer
interactions on intake upper surface significantly. Numerical results also indicate that the
nose bluntness has a significant influence on intake performance parameters.

2. COMPUTATIONAL SETUP AND PROCEDURE

    A sketch of the vehicle forebody with engine intake model with main dimensions is
shown in Fig. 2.1. All the dimensions have been non-dimensionalised using distance
from vehicle nose to intake entry, F, of the vehicle. A commercially available solver
using cell–centered finite volume technique to solve the three dimensional, compressible
Reynolds–Averaged Navier-Stokes equations code has been used. The implicit solver
with an upwind discretisation scheme for convective term and second order central
differencing scheme for diffusion terms in flow and transport equation for K-ε turbulence
model has been used. The hypersonic free stream flow is defined by specifying the
boundary conditions given in Table I. The outflow being




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                     TABLE I INFLOW BOUNDARY CONDITIONS

             Cases                       M ∞ T∞   P∞    k∞
             Two dimensional Intake with
                                         6.5 237K 830Pa 0.05
             different nose radii

predominantly supersonic, the variables are completely extrapolated from interior to the
outlet boundary. The air is assumed to be a calorically perfect gas with constant specific
heat ratio, γ = 1.4. At solid walls no-slip adiabatic boundary condition is imposed. A
rectangular computational domain was chosen           for all the cases of different nose
bluntness. The computations were performed on a 16 core cluster. The cluster was
connected to a SAN to store large amount of data generated. To monitor convergence of
the numerical solution, axial force, normal force and pitching moment plots were
monitored. The solution converged after about 25,000 iterations. An additional criterion
enforced in the current analysis required the difference between computed inflow and
outflow mass to drop to 0.5%.

        Grid sensitivity analysis confirmed that the grid resolution used is sufficient to
capture the relevant physical features. The axial force, normal force and pitching moment
obtained with different grid refinement levels were compared. Mesh level-1 (total
hexahedral cells 78,313), mesh level-2 (total hexahedral cells 2,84,355) and mesh level-3
(total hexahedral cells 8,73,635), the maximum discrepancy between the three mesh
levels was found to be less than 3%. Fig. 2.2 shows computed pressure along the vehicle
forebody lower and the intake upper surface for these




three grids. Out of these analyses, mesh level-3 was selected, and all results shown are
computed applying this resolution. To ensure the accuracy of turbulent flow solution, a
value of Y+ below 15 is maintained in the main portion of the wall flow region required
for this analysis.




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3. VALIDATION
    Suitability of the CFD solver to predict a complex high speed flow through a scramjet
inlet was evaluated by comparing the results obtained by the present solver with the
experimental results already available in the literature for flow through a dual-mode
scramjet at Mach number 4 [10]. Fig. 3.1 shows the comparison of variation of pressure
along the ramp side of dual-mode scramjet at Mach number 4. The present computational
results are also compared with computational result for the same geometry and flow
conditions [11]. The present computational results show good agreement with the already
available experimental as well as computational results.




However discrepancy on the ramp side pressure distribution can be observed in the shock
boundary layer interaction region and expansion region. The probable reason for this
discrepancy could be deficiency of the turbulence model used, assumed calorically
perfect gas condition and the differences between experimental and numerical simulation
conditions or the measurement error of the sensors. It indicates the sufficiency of grid
distribution, turbulence modeling, boundary conditions etc., being adopted in present
computations.
        Based on reasonably good comparison achieved, further computations are made
for present hypersonic intake geometry with similar grids, boundary conditions and
turbulence modeling. Fig. 3.2 shows a typical grid distribution showing the overall
computational domain with necessary boundary conditions.

4. RESULTS AND DISCUSSION
    Six two dimensional numerical simulations for different nose radii have been carried
out to obtain the static pressure distribution and skin friction distribution over different
parts of the vehicle intake. To obtain further data about boundary layer condition and
flowfield inside and outside the intake at different sections. Line probes are drawn in
vertical direction as shown in Fig. 4.1. To study the effect of nose bluntness in the sonic
line height of boundary layer two vertical line probes are drawn from one near to leading



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edge of nose at X/F= 0.24 and another at X/F= 0.48 as shown in Fig. 4.1. In addition
numerical Schlieren pictures, different contours were




     Fig. 4.1. A sketch showing naming of intake and various locations of investigations

drawn to find out flow around the vehicle nose, internal and external shock structure of
the intake. In a mixed compression intake, part of the compression takes place before the
flow enters the intake through the forebody shock and the ramp shocks. The flow is
processed through vehicle nose before it enters inside the intake. The growth of boundary
layer at the sharp leading edge of vehicle generates weak leading edge oblique shock.
This deflects the flow over vehicle body. By changing the nose from sharp to blunt, this
oblique shock changes it shape from oblique to detached bow shock. The effect of nose
bluntness by changing the nose radius is studied on external compression, internal
compression and intake performance parameters by using two dimensional computations.

4.1 Effect of nose bluntness on vehicle’s forebody lower surface pressure
    distribution
    The entropy layer is generated by the curved part of the bow shock in the vicinity of
the blunt leading edge of the vehicle, so that the streamlines passing through this curved
part of the shock form the entropy layer. The pressure of air passing along these
streamline is first increased due to stronger portion of the shock wave and after this air
expands towards the pressure of the sharp leading edge case in downstream direction. For
air travelling along the streamlines, which are further displaced from the wall thus not
passing through the curved portion of the shock, the pressure of air is only raised
equivalent to oblique shock of sharp plate. Hence the streamlines passing through the
curved portion of the shock are responsible for the overpressure observed downstream of
the vehicle leading edge as shown in Fig. 4.2. In a similar fashion press




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gradient of streamlines, passing through the curved portion of the shock represents the
entropy layer. The decay of the overpressure approximately indicates that all streamlines
of the entropy layer entered the boundary layer so that the entropy layer is swallowed
[12]. To estimate the location of entropy layer swallowing or respectively the level of the
overpressure, the blast wave theory is to be employed. The blast wave theory predicating
the overpressure extent indicates that entropy layer swallowing moves downstream with
increased leading edge radius of the vehicle nose. Therefore, small leading edge
bluntness causes the entropy layer being swallowed near the vehicle nose and for large
bluntness the entropy layer is swallowed near to the intake entry. The Fig. 4.3 shows
comparison between the flow fields with leading edge bluntness and a sharp leading edge
are sketched to demonstrate the effect of the entropy layer for different vehicle nose radii.
First the flowfield of the vehicle forebody lower surface is discussed with the help of
pressure variation as shown in Fig. 4.3. The pressure distribution on forebody lower
surface shows the increased overpressure due to increased vehicle nose radius. As nose
bluntness increases the entropy layer swallowing point location shifts from leading edge
of the vehicle nose towards the intake entry. Because of this different entropy layer
swallowing point locations for different nose radii, overpressure decays at different
locations on the forebody lower surface as shown in Fig. 4.3. The overpressure decay is
mainly determined by the leading edge bluntness and indicates that all streamlines of the
entropy layer have reached the near wall zone thus have affected the surface pressure
distribution. Since, this near surface zone is also the domain of boundary layer, the
coincidence of entropy layer swallowing and overpressure decay is somewhat reasonable
but not exact. Nevertheless, the statement allows one to consolidate the terms which have
been employed to describe the effect of leading edge bluntness on the flowfield. The
pressure variation on the vehicle forebody lower surface shows the different level of
pressure increase for different nose radii. This pressure variation is because the bow
shock strength varies with bluntness.



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4.2 Effect of nose bluntness on vehicle’s forebody lower surface skin friction
    coefficient distribution
    The streamline representing the entropy layer have passed through a stronger portion
of shock, the temperature of the air travelling along these streamline increases so that the




temperature within the boundary layer is also increased, compared to sharp leading edge
case. Due to entropy layer swallowing, the static pressure at the boundary layer edge is
the same as for the sharp leading case, so that by considering the equation of state, the
density within the boundary layer is decreased. Therefore, the boundary layer thickness is
increased and with it the sonic line height is also increased. Further, the increased
boundary layer thickness and unchanged velocity difference between the boundary layer
edge and the wall indicate a decreased velocity gradient at the wall is to be concluded.
From this decreased gradient one can conclude that wall shear stress is reduced and with
it the skin friction coefficient decreased as shown in Fig. 4.4. As a result of the
swallowed entropy layer, the temperature within the boundary layer is increased this in
turn increases the viscosity having an increasing effect on the wall shear stress. This
influence is minor as compared with the above mentioned influence of increased
boundary layer thickness thus the skin friction coefficient is decreased for the blunt
leading edge case, compared with the sharp leading edge case.

4.3 Effect of nose bluntness on sonic line height of the boundary layer on vehicle
    forebody lower surface

   For blunt nose, streamline close to the vehicle surface axis requires large defections.
As bluntness increases streamline deflection from the body axis also increases. This is
achieved by the formation of a strong near normal shock, close to the surface which
decelerates the flow to subsonic speeds, allowing it to negotiate the finite leading edge.


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Following this, the flow accelerates back to supersonic speeds. Further turning
downstream of the sonic line is achieved by Prandtl-Meyer expansion fans. Away from
the body axis, the deflection angles required are small. As a result, streamlines in this
region of the flowfield are processed by weaker oblique shock waves. This phenomenon
increases the detached bow shock angle generated by the nose as the nose bluntness
increases as shown in Fig. 4.5. The entropy layer is generated by the streamlines passing
through a curved part of the bow shock in the vicinity of the leading edge. The entropy
layer thickness increase with increase in nose bluntness. This increase in entropy layer
thickness changes the location of entropy layer swallowing point on vehicle forebody
lower surface. The stream lines passing through strong portion of the curved shock
generates entropy layer and these streamline features increased static temperature and
decreased velocity before entropy layer swallowing point and after entropy layer
swallowing point these streamlines are characterized by higher velocity and decreased
temperatures[12]. The streamline which passes through the entropy layer generates
variation in Mach number from normal to free stream flow direction over the vehicle’s
forebody lower surface as shown in Fig. 4.6 at stream wise location X/F=0.24 and at
X/F=0.48 as shown in Fig. 4.7. The streamline representing the entropy layer which
passes through entropy layer and boundary layer before entropy layer swallowing
increase the static temperature inside boundary layer and decreases the velocity of air
which is flowing through boundary layer.

        This increase in sonic speed and decrease in local velocity of air which is
travelling along with entropy layer streamline decreases the Mach number inside the
boundary layer nearer to




forebody surface. This decreased Mach number increase the sonic line height inside the
boundary layer with increase in nose bluntness at stream wise position over the vehicle’s
forebody lower surface (Fig. 4.6 and 4.7). It indicates the sonic line height increases, the


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velocity gradient inside the sonic line portion near to wall of the boundary layer
decreases. This increase in sonic line height increases the boundary layer thickness and
changes boundary layer edge conditions as compared with sharp leading edge case.

4.4 Effect of nose bluntness on pressure distribution during cow bow shock /
    boundary layer interaction on intake upper surface

   To study the cowl bow shock / boundary layer interaction on intake upper surface, a
simple flowfield model of SWBLI due to impinging shock wave on a boundary layer of a
supersonic flat




plate is shown schematically in Fig. 4.8. This figure also displays the separation of
boundary layer with corresponding separation and reattachment shock with separation
bubble.

    The upper boundary of the considered model is the streamline which would pass
through the sonic point of the undistributed boundary layer at sonic height (sh). The sonic
height is assumed to be the height of the sonic line of undistributed boundary layer
upstream of the separation point. The entropy layer generated on the vehicle forebody
lower surface influences the sonic line height of boundary layer for different nose
bluntness. This effect is discussed in earlier section in detail. How this sonic line height
of boundary layer affect the shock wave / boundary layer interaction in terms of
separation length, separation pressure and reattachment pressure is discussed in details by
observing the pressure variation on the intake upper surface and Mach contours,
numerical Schlieren pictures as shown in Fig. 4.9, 4.10 and 4.11. The flow over intake
upper surface is processed by two incident shock from the vehicle’s nose and forebody
ramp, in addition to the shock generated by the intake cowl lip. The shock wave /


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boundary layer interaction over intake upper surface is influenced by nose bluntness
effect as well as cowl shock strength. Again cowl shock strength depends on the intake
entry flow conditions which are processed through nose shock, ramp shock and their
entropy layer effect. The separation location




, reattachment location, separation length and size of separation bubble during shock
wave / boundary layer interactions depend on boundary layer thickness before
interaction, entropy layer swallowing location and type of interaction between nose shock
and cowl lip shock. The cowl shock impinges on the intake upper surface after nose and
cowl shock interaction. Thus the strength of cowl shock is influenced by which Edney
type of shock interference takes place, according to the location of nose shock
impingement on cowl lip bow shock. Present investigation is concentrated on the effect
of nose bluntness in terms of entropy layer thickness, boundary layer thickness before the
shock wave / boundary layer interactions over the intake upper surface. The effect of
increased nose radius is to increase and then decrease the separation length. The point of
reversal depends on the cowl shock strength, after nose shock and cowl shock interaction,
free stream Mach number and nose radius.

    The vehicle nose and ramp shock interact with the cowl shock at different portions of
the cowl bow shock with different nose radii as shown in Fig. 4.10 and 4.11. It may be
noted here that by carefully studying his experimental work, Edney was able to classify
six basics types of shock interference patterns, depending on the strength and relative

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direction of the intersecting shocks. The separation size increases from sharp nose case to
blunt leading edge case with radius 5.00 mm case as shown in Table II. The separation
and reattachment locations for different nose bluntness cases are located from velocity
contours and their corresponding locations are marked on pressure variation plot over the
intake upper surface as show in Fig. 4.9. This shows that for radius 10 mm the separation
length has suddenly increased as compared to radius 12.5 mm case. The probable reason
for this increase in separation length is due to the higher shock strength after nose and
cowl shock interaction and there interaction location on the intake upper surface as
shown in Fig. 4.10 and 4.11. The pressure distributions show the increased overpressure
due to the increased leading edge radius upstream of the interaction so that for the large
leading edge radii the induced pressure gradient reaches the boundary layer just upstream
of the separation point. The nose leading edge radius increase affects the pressure plateau
region, separation pressure, reattachment pressure and separation length. As nose leading
edge radius increase continuously, its effect is increase in separation and reattachments
pressures. By observing internal shock reflection pattern from numerical Schlieren and
pressure variation, it is observed that highest reattachment pressure for nose leading edge
case is of 7.5 mm. The reason for highest increase in pressure is the strongest
reattachment shock formed during shock wave boundary layer interaction. In other cases
the reattachment shock locations is before and after first expansion shoulder as shown in
Table II. This expansion fan reduces the pressure of reattachment shock for other cases.
But for nose leading case of radius 7.50 mm, the shock reattachment location is on the
expansion shoulder itself and thus it nullifies the effect of expansion fan on the
reattachment shock. This is reflected in the separation length value of 52 mm, similar to
sharp leading edge case. The effect of increase in nose leading edge radius is to shifts the
separation and reattachment locations in upstream direction as shown in Table II. This
indicates the cow bow shock wave angle increase with increase in leading edge radius
and this increased shock angle shifts the both locations is in upstream direction. If we
observe separation location for different leading edge nose radius cases, there is shifting
of separation location from sharp nose to blunt cases in upstream direction, but the
separation location of R10.00 case is in different way, as compared with other cases. The
separation location of R10.0 case should be in
   TABLE II DETAILS OF SONIC LINE HEIGHT, SEPARATION LOCATIONS,
 REATTACHMENT LOCATIONS AND SEPARATION LENGTH FOR DIFFERENT
                             NOSE RADII
           R        sh          sh       S       R    Sl (mm)
        (mm)    (X/F=0.24) (X/F=0.48) (mm)     (mm)
                    mm         mm
         0.00      0.76        0.77    2180    2230      50
         2.50      0.89        0.91    2150    2216      66
         5.00      1.26        1.27    2141    2222      81
         7.50      1.36        1.37    2113    2165      52
        10.00      1.37        1.65    2060    2159      99
        12.50      1.38        1.66    2094    2152      64




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between R7.50 and R12.50 case, but the separation location for R10.00 case is ahead of
R12.50 mm case. Reason for this could be stronger cowl shock after nose and cowl
interaction. Observations of internal shock reflection pattern indicate different nature for
different nose bluntness. This indicates the influence of nose bluntness on internal
compression also. As the shock impingement separation location moves from exit
towards the intake entry, the reattachment pressure increase which indicates that the
effect of expansion fans on reattachment shock is reduced. The analysis indicates R = 7.5
mm to be most suitable nose radius from separation length consideration.

4.5 Effect of nose bluntness on skin friction distribution during cowl bow shock /
    boundary layer interaction on intake upper surface

    The skin friction coefficient for different nose leading edge radii are shown in Fig.
4.12. The skin friction coefficient decreases continually from increase in nose leading
edge radii before the shock wave / boundary layer interactions. After the shock wave /
boundary layer interaction the distribution of skin friction coefficient on the intake upper
surface becomes very complex. This complexity is due to internal shock-shock
interactions, shock wave-expansion fan interactions and again their interaction with
boundary layer on engine cowl and intake upper surface. The skin friction coefficient
determines the velocity gradient at the wall and it defines how rapid the velocity varies
from the wall towards the boundary layer edge. Thus, it describes to some extent the
boundary layer profiles. Observing the skin friction coefficient distribution before the
shock wave / boundary layer interaction, it is to be concluded that sonic line height of the
boundary layer increases with increase in nose bluntness. This finding of the skin friction
coefficient reversal behaviour generally explains the reversal trend of the separation size
for increased nose leading bluntness. In the present study this skin friction coefficient
reversal behaviour also shows the same kind of behaviour as is observed in earlier study
as shown in Fig. 4.12. From skin friction coefficient distributions it can be concluded that
skin friction coefficient behaviour before interaction is different and after interaction it is
different, not following the same trend




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which was there before interaction with increase in nose bluntness. The skin friction
coefficient is observed to be highest in nose leading edge case of 7.5 mm. The pressure
variation over intake upper wall for this case had shown the highest pressure increase at
the reattachment shock. This indicates that skin friction coefficient after shock wave /
boundary layer depends on the reattachment shock strength. As the velocity before
separation point is reduced to almost zero, the skin friction coefficient is observed to be
minimum for all separation locations measured from velocity contours. It indicates that
the measured separations and reattachment locations are correct. The observed skin
friction coefficient behaviour at reattachment point is different than the expected one.
The skin friction coefficient should be higher after reattachment locations, but observed
trend for skin friction coefficient is different. The reason for this could be that the
reattachment shocks interact with the expansion fans generated from the first expansion
shoulder. This phenomenon may be changing the skin friction coefficient behaviour at
reattachment locations.

4.6 Effect of nose bluntness on pressure distribution and skin friction distribution
    on engine’s cowl surface

The distribution of computed pressure ratio variation and skin friction coefficient
variation on the engine’s cowl surface is shown in Fig. 4.13 and 4.14. As nose bluntness
increases the initial pressure on the engine cowl surface is slightly increased. Observation
of pressure distribution indicates the impingement of separation and reattachment shock
formed during the shock wave / boundary layer interactions impinges on the cowl surface
at different locations. This produces different pressure rise at different locations on the
engine cowl surface. In 12.5 mm nose bluntness case the effect of expansion waves is not
seen on the cowl pressure distribution. The reason for the same is that, the separation and
reattachment shocks locations are before the expansion shoulders. The expansion waves
formed at expansion shoulder interacts with separation and reattachment shock at some


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distance from the cowl surface and it’s effect is not seen on cowl pressure distribution,
but this interaction reduces the pressure rise on the cowl surface as compared to other
bluntness cases (Fig. 4.13). Nose bluntness case of 7.5 mm produces pressure rise on the
engine cowl in upstream direction as compared to other cases. The computed skin friction
coefficient distribution on engine’s cowl surface for different nose radii is shown in Fig.
4.14. Initial skin friction coefficient distribution shows that it is not following the same
trend as in case of increasing nose radius in the magnitude of skin friction coefficient.
This skin friction coefficient depends on the cowl and nose shock interaction among them
and after interaction, type of shock interference formed. Initial skin friction coefficient
distribution shows the skin friction coefficient reduction for nose bluntness case of sharp,
2.5 mm and 5.0 mm as compared with other cases. Observation of skin friction
coefficient shows the reduction of skin friction coefficient at reflected shock
impingement locations and, where the expansion waves reach to engine cowl surface
their increase is observed.




4.7 Effect of Nose Bluntness on Performance Parameters
    Hypersonic intake performance can be assessed in form of various parameters.
Typical performance parameters are the mass flow rate, total pressure recovery ( ) and
adiabatic kinetic energy efficiency              . When isentropic expansion is considered,
the adiabatic kinetic efficiency is the ratio of kinetic energy of the decelerated flow to the
kinetic energy of the undistributed flow. Total pressure recovery is defined as total
pressure ratio between the exit and entry plane of the intake. Mass flow rate, total
pressure recovery, intake exit Mach number and adiabatic efficiency parameters of the
inlet obtained from the numerical simulations are shown in Fig. 4.16. All the values are
by averaging of area of the respective variables at indicated locations. Nose bluntness

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affects all the performance parameters of the intake because the flow is processed
through nose shock and ramp shock before entering the inlet. Mass flow rate increases for
sharp nose case to blunt case of 2.5 mm radius and further the decrease in mass flow rate
is observed from 2.5 mm radius to 12.5 mm radius nose. This decrease is due to different
location of nose shock interaction with cowl shock. Total pressure recovery is slightly
increased from sharp case to 2.5 mm nose radius case. Further decrease in pressure
recovery is observed from 2.5 mm nose radius case to 12.5 mm case. The causes for the
reduction may be stronger shock wave / boundary layer interaction inside the intake. The
intake exit Mach number is first reduced from sharp case to 10.0 mm nose radius case
and 12.5 mm case Mach number is further increased. This indicates that the Mach
number is dependent on the cowl shock strength which impinges on the intake upper
surface and its further reflection inside the intake. Adiabatic kinetic energy efficiency
being dependent on exit velocity follows the same trend as exit Mach number.

5. SUMMARY AND CONCLUSIONS

    Numerical simulations of turbulent, compressible, 2-D viscous flow in the hypersonic
intake for various nose radii are presented. The numerical methodology has been
validated by simulating the external and internal flow of dual mode scramjet intake and
comparing with experimental data.
    As bluntness increases the pressure distribution over vehicle forebody lower surface,
intake upper surface and engine cowl increases. Increase in nose bluntness decreases the
skin friction coefficient over vehicle forebody lower surface, intake upper surface. Sonic
line height of the boundary layer increases over the vehicle forebody lower and intake
upper surface with increase in nose radii. Nose bow shock angle increases with increase
in nose radius. Separation length trends are not fixed with increase in nose bluntness.
Separation length increase is observed up to the nose radius case of 5.00 mm and after
this again separation length is decreased for 7.5 mm case. Both separation and
reattachment locations get shifted in upstream direction with increase in nose radius,
except for 12.5 mm nose radius case. Mass flow rate and total pressure recovery was
found to increase for 2.5 mm radius case and for remaining cases it continuously
decreases with bluntness. Intake exit Mach number and kinetic energy efficiency was
observed to be higher for 12.5 mm nose radius case.

6. REFERENCES
[1] Holland.S.D and Murphy.K.J. 1993. “An experimental parametric study of geometric,
Reynolds number and specific heats effects in three dimensional sidewall compression
Scramjet intakes at Mach 6”. AIAA Paper 93-0740.
[2] Holland Scott D 1995 “Computational parametric study of sidewall-compression
scramjet inlet performance at Mach 10”. NASA Technical Memorandum 4411.
[3]     Goonko.Y.P., Latypov.A.F. Mazhui.I.I. Kharitonov.A.M and Yarosalavtsev.
2003. “Structure of flow over a hypersonic intake with sidewall wedges”. AIAA Journal,
Vol.41, No.3, pp 436-447.
[4]     Mahapatra. D and Jagadeesh. G. 2008. “Shock tunnel studies on cowl/ramp
interactions in generic scramjet intake”. Journal of Aerospace Engineering, Vol.222, pp
1183-1191.


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[5] Sivkumar.R and Babu.V. 2006. Numerical simulation of flow in a 3-D supersonic at
higher Mach number. Defence Science Journal, Vol.56, No.4, pp 465-476.
[6] J-H Liang, X-Q Fan, Y Wang and W-D Liu. 2008. “Performance enhancement of
three –dimensional hypersonic inlet with sidewall compression”. Proc. I Mech Vol.222
Part G: J. Aerospace Engineering, pp1121- 1219.
[7] Sanator, R.J, Boccio, J.L and Shamshins, D. 1968           “Effects of Bluntness on
Hypersonic Two-Dimensional Inlet type Flows.” NASA CR-1145.
[8] Townsend, J.C “The Effects of Leading Edge Bluntness and Ramp Deflection Angle
on Laminar Boundary Layer Separation in Hypersonic Flow” NASA TN-D 3290.
[9] Daniel Arnal and Jean Delery May 2004: SWBLI, NATO.
[10] Saied. Emani, Carl. A. Texler, Aaron. H. Auslender and John. P. Weidner. 1995
“Experimental investigation of inlet-isolator combustor for a dual mode scramjet at a
Mach number of 4.” NASA Technical paper 3501.
[11] Chang. J, D. Yu, W. Bao, Z. Xie and Y. Fan. 2008 “A CFD assessments of
classifications for hypersonic inlet start/unstart phenomena” The aeronautical Journal,
volume 113, No 1142.
[12] Hirschel M.S. 2005. “Basics of Aerothermodynamics” Springler Verlag, Berlin.
[13] Amarjit Singh and J. L. Stollery. 1995. “Experimental investigation of hypersonic
flow over a wing-body combination”, AIAA-95-6083, Presented at AIAA sixth
international aerospace planes and hypersonic technologies conference, Chattanooga, TN,
USA.
[14] Jaware. V. B, Amarjit Singh and Joshi. S. V. 2009. “Numerical simulation of
internal flowfield in a 2D and a 3D hypersonic mixed inline intake at Mach 6.5.”,
International Conference on “Latest Trends in Simulation Modelling and Analysis
(COSMA2009)”, Proceedings of National Institute of Technology Calicut, Calicut,
Kerala, India, VITF 76, pp 422-425.
[15] Jaware. V. B, Amarjit Singh and Joshi. S.V. 2009. “A computational study of effect
of intake fence on the performance of hypersonic vehicle inlet”. Proceedings of Thirty
Sixth National Conference On “Fluid Mechanics and Fluid Power (NCFMFP2009)”,
Government College of Engineering, Shivaji Nagar, Pune, India , pp237-243.
[16] Jaware. V. B, Amarjit Singh and Joshi.S.V. 2009 “A computational study of internal
flowfield investigation of a hypersonic vehicle inlet at mach 6.5.” Proceedings of
National Conference on “Modelling and Simulation (NCMS2009)”, Defence Institute of
Technology, Girinagar, Pune, India,
[17] Mahapatra. D and Jagadeesh. G. 2009. “Studies on unsteady shock interactions
near a generic scramjet intake.” AIAA Journal, Vol. 47, No.9, pp 2223-2231.
[18] Delery, J. M. 1999. “Shock phenomena in high speed aerodynamics: Still a source of
major concern”. The Aeronautical Journal, vol.103, Paper No.1019, pp 19-34.
19]     Jaware. V. B, Amarjit Singh and Joshi S.V. 2010. “Numerical simulation of
internal flowfield and performance evaluation of a 3D mixed inline intake with fence”,
Proceedings of International Conference on Advances in Mechanical Engineering
(ICAME-2010))” held at S. V. National Institute of Technology (SV NIT), India, pp 332-
326.
[20] Jaware. V. B, Amarjit Singh and Joshi S.V. 2011. “Computational study of the effect
of fence on the external and internal compression in a mixed inline intake” International
Review of aerospace Engineering (IRSE) Vol. 4, No.2,pp 35-42.


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[21] Jaware. V. B, Amarjit Singh and Joshi S.V. 2011 “Numerical simulation of external
and internal flowfield of a 3D mixed inline intake with strake” Proceeding of National
conference on ‘Advances in Engineering, Management and General sciences’ , Pimpri-

7. Authors Information

                Prof. VB Jaware
                Prof. Jaware is presently working as Assistant Professor in Mechanical
                Engineering, Department at Rajarshi Shahu College of Engineering, Tahawade,
                Pune, Maharashtra, India. He has obtained his BE (Mechanical Engineering) in
                1995 from Marathwada Institute of Technology, Aurangabad                and ME
                (Mechanical Heat Power Engineering) in 2003 from Vishwakarma Institute of
Technology, Pune. At present he is working for his doctoral studies at University of Pune.
        His current research area is hypersonic internal and external flows. He has published
twenty three research papers in International and National Journals and conferences. He has co-
authored one text book for Mechanical Engineering UG course of University of Pune. Prof.
Jaware is a life member of Indian Society of Technical Education (ISTE), India.

Dr. Amarjit Singh

                  Dr. Singh is presently working as Scientist ‘G’ and Associate Director at
                  Terminal Ballistic Research Laboratory, Chandigarh, India. He has obtained his
                  BE (Aeronautical Engg) from Punjab Engineering College Chandigarh in 1980,
                  M Tech (Mechanical) (Guided Missiles) from University of Pune in 1985 and
                  PhD (Aerospace) from Cranfield University, UK in 1996. Some of his specific
                  contributions are: Setting up test and research facilities such as subsonic and
                  supersonic wind tunnels, campus wide network CROWN, and advanced
computing facilities at DIAT Pune. He has worked on a number of projects sponsored by DRDO.
  His current areas of interest include: High-speed internal and external flows, shock wave
boundary layer interactions, aircraft design, and computational fluid dynamics. He has published
40 research papers in international and national journals and conferences. Dr. Singh is a life
member of Aeronautical society of India and Astronautical Society of India.

                  Dr. Satishchandra Joshi

                  Dr. Joshi is presently working as Principal at D.Y.Patil College of Engineering,
                  Pune, Maharashtra, India. He obtained his B Tech (Mechanical Engineering)
                  from Indian Institute of Technology Madras in 1978, M. Engg (Energy
                  Technology) from Asian Institute of Technology, Bangkok, Thailand in 1985
                  and PhD (Energy systems Engineering) from Indian Institute of Technology,
Bombay in 2005. Some of his specific contributions are: Setting up test and research facilities in
IC Engines alternative fuels, Centre of Energy and Environmental Studies at VIT Pune. He has
worked on a number of research and consultancy projects sponsored by University of Pune and
industry in USA, and India.
         His current areas of interest include; Alternative fuels, pollution, Stirling engines, Energy
Conservation and Renewable Energy resources like Solar and Biomass. He has published 24
research papers in international and national journals and conferences.Dr. Satishchandra Joshi is a
life member of ISTE (Indian Society of Technical Education), ISHMT (International Society of
Heat and Mass Transfer), and Solar Energy Society of India (SESI). He is a Fellow of the
Institution of Engineers (India).


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