International Journal of Mechanical Engineering and Technology … · 2014. 3. 4. · Suitability of the CFD solver to predict a complex high speed flow through a scramjet inlet was

International Journal of Mechanical Engineering and Technology (IJMET), ISSN 0976 –

6340(Print), ISSN 0976 – 6359(Online) Volume 3, Issue 1, January- April (2012), © IAEME

161

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

INTERNATIONAL JOURNAL OF MECHANICAL

ENGINEERING AND TECHNOLOGY (IJMET)

ISSN 0976 – 6340 (Print) ISSN 0976 – 6359 (Online)

Volume 3, Issue 1, January- April (2012), pp. 161-178

© IAEME: www.iaeme.com/ijmet.html

Journal Impact Factor (2011): 1.2083 (Calculated by GISI)

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IJMET

© I A E M E



162

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

different nose radii 6.5 237K 830Pa 0.05

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

(mm)

sh

(X/F=0.24)

mm

sh

(X/F=0.48)

mm

S

(mm)

R

(mm)

Sl (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,

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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).

Documents

International Journal of Mechanical Engineering and Technology … · 2014. 3. 4. · Suitability of the CFD solver to predict a complex high speed flow through a scramjet inlet was