Prediction of Supersonic Axisymmetric Air Intake Performance through

Numerical Simulation

A PROJECT WORK

Submitted in fulfillment of the award of Degree of Bachelor of

Technology in Aeronautical Engineering

Submitted by

M.Akhilesh (12951A21C2)

Ch. Bapi Raju (12951A21C4)

Y.Ganesh Babu (12951A21D0)

T.Rohith Kumar (12951A21D8)

Under the Supervision of

PROF. K. BHARADWAJAN

Department of Aeronautical Engineering

INSTITUTE OF AERONAUTICAL ENGINEERINGDUNDIGAL – 500 043, HYDERABAD, TELANGANA STATE

April, 2016

INSTITUTE OF AERONAUTICAL ENGINEERINGDUNDIGAL – 500 043, HYDERABAD, TELANGANA STATE

Department of Aeronautical Engineering

CERTIFICATE

This is to certify that the work embodies in this dissertation entitled ‘Prediction Of Supersonic Axisymmetric Air Intake Performance Through Numerical Simulation’ being submitted by M.AKHILESH (12951A21C2), CH.BAPIRAJU (12951A21C4), Y.GANESHBABU (12951A21D0), T.ROHITH KUMAR (12951A21D8) for partial fulfillment of the requirement for the award of Bachelor of Technology in Aeronautical Engineering discipline to Institute of Aeronautical Engineering, Dundigal, Hyderabad, Telangana State, during the academic year 2014-2015 is a record of Bonafide piece of work, undertaken by him/her in the supervision of the undersigned. Approved and Supervised by

(Prof. K.Bharadwajan) Aeronautical Engineering

Forwarded by

Dr. P. K. Dash Dr. D. GovardhanDean Academics Aeronautical Engineering IARE, Hyderabad IARE, Hyderabad

i

ii

INSTITUTE OF AERONAUTICAL ENGINEERING

DUNDIGAL – 500 043, HYDERABAD, TELANGANA STATE

Department of Aeronautical Engineering

DECLARATION

We M.AKHILESH, CH.BAPIRAJU, Y.GANESH BABU, T.ROHITH KUMAR’,

are students of Bachelor of Technology in Aeronautical Engineering’,

session: 2012 - 16, Institute of Aeronautical Engineering, Dundigal, Hyderabad,

Telangana State, hereby declare that the work presented in this project work

entitled ‘Prediction of Supersonic Axisymmetric Air Intake Performance

Through Numerical Simulation’ is the outcome of our own bonafide work and is

correct to the best of our knowledge and this work has been undertaken taking care

of engineering ethics. It contains no material previously published or written by

another person nor material which has been accepted for the award of any other

degree or diploma of the university or other institute of higher learning, except

where due acknowledgment has been made in the text.

M.Akhilesh (12951A21C2)

Ch. Bapi Raju (12951A21C4)

Y.Ganesh Babu (12951A21D0)

T.Rohith Kumar (12951A21D8)

iii

INSTITUTE OF AERONAUTICAL ENGINEERINGDUNDIGAL-500043, HYDERABAD, TELANGANA STATE

Department of Aeronautical Engineering

ACKNOWLEDGEMENT

We earnestly take the responsibility to acknowledge following distinguished personalities who

graciously allowed us to carry out my our project work successfully.

We would like to express our sincere thanks to Dr. Amalesh Barai, Principal,

Dr.D.Govardhan, Head of the department, Aeronautical Engineering,

Institute of Aeronautical Engineering, Hyderabad for their consistent guidance

and encouragement which lead to the successful completion of the project work. We express our gratitude to our college Institute of Aeronautical Engineering, Hyderabad for

providing means of attaining of my most cherished goals. We profoundly thank my project

guide Prof. K. BHARADWAJAN who has been an excellent guide and also a great source of

inspiration to my work in carrying out my industrial oriented major project.

The satisfaction and euphoria that accompany the successful completion of the task

would be great but incomplete without the mention of the people who made it possible with their

constant guidance and encouragement crowns all the effort with success. In this context we

would like to thank all other staff members of the department who have extended their timely

help and eased our work.

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ABSTRACT

To predict performance evaluation of an supersonic axisymmetric air inlet through

numericalsimulation.To increase the individual component efficiency by overcoming the

problems at intake. First, the dimensions and physical parameters of the required inlet are

taken. Then generate the axisymmetric air intake model using GAMBIT, applying

relevant boundary conditions (for optimum mach number in supersonic flight conditions),

quantification of total drag experienced, analyze the aerodynamic performance and

individual component efficiency.

Perform numerical computations using FLUENT and obtain dynamic flow field

characteristics like velocity, total pressure distribution, mass flow ratio, boundary layer

seperation. Validation of obtaned data frm numercal results with the experimental data

provided are the desired outcomes of the proposed project

v

TABLES OF CONTENTS

S.No. Description Page. No.

1 Chapter 1 – Introduction 1

1.1 Over view of the project 3

2 Chapter 2 – literature review 4

2.1 Supersonic inlets 5

2.2 Ramjet stationing 6

2.3 Types of supersonic inlets 7

2.4 Axisymmetric inlets 7

2.5 Two dimensional 8

2.6 Types of compressions 8

3 Chapter 3 – supersonic Axisymmetric inlet 10

3.1 Introduction 10

3.2 Geometry 11

3.3 Modes of operation 11

3.4 performance 14

3.5 Influencing parameters 15

3.6 Relevant technical issues 16

4 Chapter 4 – problem definition 19

4.1 Flow separation 19

4.2 Pressure recovery 20

4.3 Inlet performance 21

4.4 Boundary layer bleed 22

4.5 Bleed system functions 23

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S.No. Description Page. No.

5 Chapter 5 – Introduction 24

5.1 Gambit 24

5.2 Create Geometry In Gambit 25

5.3 Mesh geometry in gambit 29

5.4 Specify Boundary Types In GAMBIT 33

5.5 Fluent 35

5.6 Initial Settings 37

5.7 Boundary Conditions 39

5.8 Operating Conditions 39

5.9 Solution Methods 40

5.10 Solution Controls 41

5.11 Monitors 42

5.12 Solution Initializations 42

5.13 Run Calculations 43

5.14 Results and Comparisons 43

6 Chapter 6 – Conclusion 57

7 Chapter 6 – References 58

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LIST OF FIGURES

Fig. No. Description Page No.2.1 Schematic Sketch of Air Breathing Engine 4

2.2 Geometry of Supersonic Inlet 5

2.3 Ramjet 6

2.4 T-S Diagram Of Ramjet 6

2.5 Axisymmetric Inlet and Its Parts 7

2.6 Rectangular Supersonic Inlet 8

2.7 External Compression Inlet 8

2.8 Internal Compression Inlet 9

2.9 Mixed Compression Inlet 9

3.1 Supersonic Axisymmetric 10

3.2 Critical Inlet Operation 12

3.3 Subcritical Inlet Operation 12

3.4 Supercritical Inlet Operation 13

3.5 Boundary Layer Separation with Zones 16

3.6 Oblique Shock Diffuser 16

3.7 Phases Of Buzzing 17

3.8 Supersonic Inlet Relevant Technical Issues 18

4.1 Boundary Layer Separation with Zones 19

4.2 Geometry and Operation Of The SR-71 Mixed Compression

Inlet

20

4.3 Intake Shock System with Intake Critical (Ideal) 20

4.4 Starting Of an Intake 21

4.5 Represents Scramjet Flow Path With Boundary Layer Bleed And Corresponding Station Numbering

22

4.6 Spike Bleed Exit Louver 23

5.1 Gambit Topology 25

5.2 Create real vertex 26

5.4 Creating Edges 28

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5.5 Creating Faces 29

5.6 Mesh edges 30

5.7 Edge mesh 31

5.8 Complete edge mesh 31

5.9 Mesh faces 32

5.10 Face mesh 33

5.11 Specify boundary types 34

5.12 Export mesh file 35

5.13 Gambit- Fluent 36

5.14 Fluent launcher 37

5.15 General 37

5.16 Models 38

5.17 Materials 38

5.18 Pressure Far Field 39

5.19 Operating Conditions 39

5.20 Reference Values 40

5.21 Solution Methods 41

5.22 Solution Controls 41

5.23 Residual Monitors 42

5.24 Solution Initialization 42

5.25 Run Calculation 43

5.26 Scaled Residuals with Slot 43

5.27 Scaled Residuals without Slot 44

5.28(a) Temperature Contour with slot 44

5.28(b) Temperature Contour without slot 45

5.29(a) Static Temperature Contour with slot 45

5.29(b) Static Temperature Contour without slot 46

5.30(a) Static Pressure Contour with slot 46

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5.30(b) Static Pressure Contour without slot 47

5.31(a) Velocity Contour with slot 47

5.31(b) Velocity Contour without slot 48

5.32(a) Enthalpy Contour with slot 48

5.33(a) Density Contour with slot 49

5.33(b) Density Contour without slot 49

5.34(a) Plot of Total Pressure with Slot 50

5.34(b) Plot of Total Pressure without Slot 50

5.35(a) Plot of Static Pressure with Slot 51

5.35(b) Plot of Static Pressure without Slot 51

5.36(a) Plot of Static Temperature with Slot 52

5.36(b) Plot of Static Temperature without Slot 52

5.37(a) Plot of Total Temperature with Slot 53

5.37(b) Plot of Total Temperature without Slot 53

5.38(a) Plot of velocity with Slot 54

5.38(b) Plot of velocity without Slot 54

5.39(a) Plot of Enthalpy with Slot 55

5.39(b) Plot of Enthalpy without Slot 55

5.40(a) Plot of Density with Slot 56

5.40(b) Plot of Density without Slot 56

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LIST OF SYMBOLS AND ACRONYMS

° - Degrees

m - Mach number

SR - Strategic recoinnance

Po,f - Stagnation pressure at engine

face

P0, ∞ - Stagnation pressure in free

stream

Re - Reynolds Number

mR - mass flow ratio,

AOA - Angle of Attack

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LIST OF TABLES

Table No. Description Page No.3.1 Modes of operation 13

5.1 Position of vertex points 25

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1

Report OverviewChapter 1Project Overview

CHAPTER 1

Introduction

1.1 Project Overview

An inlet for a supersonic aircraft, on the other hand, has a relatively sharp lip. The inlet lip is sharpened to minimize the performance losses from shock waves that occur during supersonic flight. For a supersonic aircraft, the inlet must slow the flow down to subsonic speeds before the air reaches the compressor. Some supersonic inlets, like the one at the upper right, use a central cone to shock the flow down to subsonic speeds. Other inlets, like the one shown at the lower left, use flat hinged plates to generate the compression shocks, with the resulting inlet geometry having a rectangular cross section.

This variable geometry inlet is used on the F-14 and F-15 fighter aircraft. More exotic inlet shapes are used on some aircraft for a variety of reasons.

An inlet must operate efficiently over the entire flight envelope of the aircraft. At very low aircraft speeds, or when just sitting on the runway, free stream air is pulled into the engine by the compressor. In England, inlets are called intakes, which is a more accurate description of their function at low aircraft speeds.

At high speeds, a good inlet will allow the aircraft to maneuver to high angles of attack and sideslip without disrupting flow to the compressor. Because the inlet is so important to overall aircraft operation, it is usually designed and tested by the airframe company, not the engine manufacturer. But because inlet operation is so important to engine performance, all engine manufacturers also employ inlet aerodynamicists. The amount of disruption of the flow is characterized by a numerical inlet distortion index. Different airframes use different indices, but all of the indices are based on ratios of the local variation of pressure to the average pressure at the compressor face.

The ratio of the average total pressure at the compressor face to the free stream total pressure is called the total pressure recovery. Pressure recovery is another inlet performance index; the higher the value, the better the inlet. If the airflow demanded by the engine is much less than the airflow that can be captured by the inlet, then the difference in airflow is spilled around the inlet.

2

The airflow mismatch can produce spillage drag on the aircraft. The aim of project is to provide literature survey in the relevant state-of-the-art.

3

Aim

CHAPTER 2

LITERATURE SURVEY

Air-Breathing Inlets Most modern passenger and military aircraft are powered by gas turbine engines, which are also called jet engines. There are several different types of gas turbine engines, but all turbine engines have some parts in common. All turbine engines have an inlet to bring free stream air into the engine. The inlet sits upstream of the compressor and, while the inlet does no work on the flow, inlet performance has a strong influence on engine net thrust. As shown in the figures above, inlets come in a variety of shapes and sizes with the specifics usually dictated by the speed of the aircraft. The inlet interchanges the organized kinetic and random thermal energies of the gas in an essentially adiabatic process. The perfect inlet would thus correspond to an isentropic process. The primary purpose of the inlet is to bring the air required by the engine from free stream conditions with minimum total pressure loss.[1]

Fig.2.1 Schematic Sketch of Air breathing Engine [1].

The performance of an inlet is related to the following characteristics:

High total pressure ratio Controllable flow matching of requirements Good uniformity of flow Low installation drag Good starting and stability Low signatures (acoustics, radar, etc.) Minimum weight and cost Reliability goals

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2.1 Supersonic Inlets

In the contrary, an inlet for a supersonic aircraft has a relatively sharp lip. The inlet lip is sharpened to minimize the performance losses from shock waves that occur during supersonic flight. For a supersonic aircraft, the inlet must slow the flow down to subsonic speeds before the air reaches the compressor. Some supersonic inlets, like the one at the upper right, use a central cone to shock the flow down to subsonic speeds. Other inlets, like the one shown at the lower left, use flat hinged plates to generate the compression shocks, with the resulting inlet geometry having a rectangular cross section. This variable geometry inlet is used on the F-14 and F-15 fighter aircraft. More exotic inlet shapes are used on some aircraft for a variety of reasons. The inlets of the Mach 3+ SR-71 aircraft are specially designed to allow cruising flight at high speed [3].

Fig.2.2: Geometry of supersonic inlet [3].

2.2.1 Principle:

Flow in a supersonic inlet is much more problematic than the flow in a subsonic inlet. The big difference is that flow moving faster than the speed of sound has no knowledge of what is ahead of it. The result of this makes the design of a supersonic inlet similar to laying out the course for a slalom skier, only in this case the skier is blind.

5

2.2 RAMJET Stationing:

Fig.2.3: Ramjet [Ref 4]

P-v DIAGRAM T-S DIAGRAM T-P DIAGRAM

Fig.2.4 T-S diagram of ramjet [Ref 3]

6

2.3 Types of Supersonic Inlets

Axisymmetric or two-dimensional • Axisymmetric: central cone for shock fixture • Two-dimensional: rectangular cross-section

Variable or fixed geometry • Variable: the central cone may be movable or in a rectangular intake, one of the walls may be adjustable • Fixed: Geometry is fixed.

2.4 Axisymmetric Inlet

According to Bendot et al. (1984) the performance of an axisymmetric inlet deteriorates rapidly with an increase in angles of attack for two-dimensional and chin inlet configurations an improvement in performance with increasing positive (nose-up) angles of attack is encountered. The role of engine inlet is very vital for all air-breathing propulsion systems; especially for high supersonic in which all the compression process is performed by the inlet.

Fig.2.5 Axisymmetric Inlet and Its Parts [Ref1].

7

2.5 Two Dimensional Inlets:

Fig.2.6 Rectangular supersonic inlet [Ref 2].

2.6 Types of compressions:

2.6.1 External Compression:

The supersonic portion of the compression is done externally, ahead of the cowl lip. An advantage of a two-wedge external compression inlet is that the construction is relativelysimple and the optimum configuration is normally determined experimentally in a wind-tunnel.

Fig.2.7 External compression inlet [Ref 6]

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2.6.2 Internal Compression:Supersonic compression is done internally aft of the cowl lip. This type of inlet has a low drag value because external deflection and disturbance of the entering flow are prevented.

Fig.2.8 Internal Compression Inlet [Ref 6]

2.6.3 Mixed Compression: Supersonic compression is done both externally, forward of the cowl lip and internally, aft of the cowl lip. The advantage of a better stagnation pressure recovery is cancelled out to a great extent by the need to control Bow separation in the inlet duct. The flow separation is caused by the interaction of the internalShock systems (oblique and normal shocks) with the boundary layer.

Fig.2.9 Mixed compression inlet [Ref 6]

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CHAPTER 3

SUPERSONIC AXISYMMETRIC INLET

3.1 Introduction

The supersonic inlet consists of a spike (center-body or fore-body) and an integrated duct. The

initial compression is done by the spike. When designing a supersonic inlet, an increase in the

flight Mach number requires the increase in the number of oblique shocks in order to save total

pressure recovery. Therefore, the principle of staging a supersonic compression like the

Oswatitsch principle [Ref 3] is used to reduce the inlet losses in a most efficient manner.

Fig.3.1 Supersonic Axisymmetric Inlet

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3.2 Geometry An inlet generally consists of the following components:

Ramp: An intake ramp is a rectangular, plate-like device within the air intake of a jet engine, designed to generate a shock wave to aid the inlet compression process at supersonic speeds. The ramp sits at an acute angle to deflect the intake air from the longitudinal direction

Throat: A throat is minimum area between the centre body and cowl. Cowl: Cowl is upper and lower part of the engine. Subsonic diffuser: It is a divergent profile to increase the pressure of the flow. [3]

3.3 Modes Of Operation:

3.3.1 Critical Operation

The condition when the inlet can accept the mass flow of air required to position the terminal shock just inside the cowl lip is called critical inlet operation.

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Modes of operationCriticalSubcriticalSupercritical

Fig.3.2 Critical Inlet Operation [Ref 4].

3.3.2 Subcritical Operation

The condition when the inlet is not matched to the engine, due to which the normal shock moves upstream and stays in front of cowl lip, is called as sub-critical operation

Fig.3.3 Subcritical Inlet Operation [Ref 4].

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3.3.3 Supercritical Operation:

The condition when the inlet cannot capture the mass flow required by the engine and the terminal shock is sucked into the diffuser is called super – critical operation

Fig.3.4 Supercritical inlet operation [Ref 4].

Modes of operation

Axisymmetric inlet SketchPosition of

Normal shock wave

Sub-Critical

The normal shock moves upstream and stays in front

of cowl lip

CriticalThe terminal

shock just inside the cowl lip

Super-CriticalThe terminal

shock is sucked into the diffuser

3.1 Modes of operation

13

3.4 Performance:

The performance of an inlet is characterized by three important parameters, namely pressure recovery, mass flow ratio and boundary layer bleed. Each one of the above-mentioned has an effect on the overall or total drag figure of an inlet.

3.4.1 Pressure Recovery

The inlet total pressure recovery is defined by Seddon (1988:5) as the ratio of the mean totalpressure at the engine face to the total pressure available in the free stream, infinitely farupstream of the inlet.

Po,f = Stagnation pressure at engine faceP0, ∞ = Stagnation pressure in free stream

The total pressure recovery is an indication of the maximum pressure available in the combustion chamber and resultant thrust that can be developed. With an increase in free stream (flight) Mach numbers, high shock losses and shock wave boundary layer interactions cause a decrease in total pressure recovery. Methods to remedy these high shock losses, ego variable compression surface where oblique shocks are prevented from entering into the inlet, can result in an increase in drag. This is as a result of the greater angle through which the flow is turned before it enters the inlet opening.

3.4.2 Mass Flow Ratio

The mass flow ratio is defined by Gregoriou (1985:7} as the ratio of air mass flow at inlet entryto air mass flow at free stream conditions.

Where,

14

This mass flow ratio, mR, can also be expressed as an area ratio since the equation for massflow (Hall 1977:12) can be written as follows:

3.5 Influencing Parameters:

3.5.1 Total Pressure Ratio: The ratio total pressure at inlet is given by the ratio of total pressure leaving at stage 2 and the total pressure entering at stage 0.

Πd = Pt2/Pt0

3.5.2 Total Temperature Ratio:

The ratio total temperature at inlet is given by the ratio of total temperature leaving at stage 2 and the total temperature entering at stage 0. Ʈd = Tt2/Tt0

3.5.3 Isentropic Diffuser Efficiency

3.5.4 Area Mach Relation

15

3.6 Problems

3.6.1 Boundary Layer Separation

Separation of the external flow in zone 1 may result from local high velocities and subsequent deceleration over the outer surface and it leads to high nacelle drag.

Separation on the internal surfaces may take place in either zone 2 or zone 3, depending on the geometry of the duct and the operating conditions.

Zone 3 may be the scene of quite large adverse pressure gradients since the flow accelerates around the nose of the center body and then decelerates as the curvature decreases.

Fig.3.5 Boundary Layer Separation with Zones [Ref 7]

3.6.2 External Deceleration

The simplest and most practical external deceleration in simple oblique shock wave is in some cases, a series of oblique shock waves.

Fig.3.6 Oblique shock diffuser [Ref 7]

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3.6.3 Buzzing:

Buzz is an airflow instability caused by the shock waves rapidly being alternately swallowed and expelled at the inlet of the duct and occurs in supersonic intakes at subcritical operations. It starts when the aircraft begins to fly at or near the speed of sound. At these speeds sonic shock waves are developed that if not controlled will give high duct loss in pressure and airflow and will set up vibrating conditions in the inlet duct, called inlet Buzz.[8].

Fig.3.7 Phases of buzzing [Ref 8].

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Region of interest

Fig.3.8 Supersonic Inlet Relevant Technical Issues

We will focus on the problem spike bleed and we will proceed to how to solve it efficiently.

CHAPTER 418

Supersonic inlet relevant technical issues

Problem Definition

The boundary layer on the cone is stretched as it moves up the cone preventing flow separation , but for the internal compression and the subsonic compression the boundary layer still tends to separate and usually is sucked through tiny holes  in the wall. As a side note on the aero spike engine the boundary layer gets thicker towards the end of the cone as needed for the greater speed difference between the air molecules just on the surface of the cone and the fully accelerated stream of air.

Fig.4.1 Boundary Layer Separation with Zones [Ref 7]

OBJECTIVES

4.1 FLOW SEPERATION

Supersonic flow over spiked cylinders nonsteady regimes can occur in which a separation zone is periodically generated at the spike. The physical pattern of flow with separation zone fluctuations and have determined the boundaries of existence of the nonsteady regime as a function ratio between the spike length and diameter of the cylinder.

19

Fig.4.2 Details of the Geometry and Operation of the SR-71 Mixed Compression InletBy J. Thomas Anderson Technical Fellow Emeritus Lockheed Martin [6].

4.2 Pressure Recovery

Pressure recovery charecteristics of conical spike inlet with a fixed area bypass located in the top or bottom of the diffuser are presented for fligh mach numbers of 1.6, 1.8, and 2.0 for angle of

attack 0 degrees to 9 degrees.Top or bottom location of bypass did not have significant effects on diffuser pressure recovery,bypass mass flow ratio,or drag coefficient over the range of angle of

attack,flight mach numbers and stable engine mass flow are investigated.

Fig.4.3 intake shock system with intake critical

At a flight Mach number of 2.0 the discharge of 14 percent of the critical mass flow of the inlet by means of by pass increased the drag only one-fifth of the additive drag that would result for equivalent spillage behind an inlet normal shock without significant reductions in pressure recovery.

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4.3 Inlet Performance

Supersonic diffusers are characterized by the presence of shocks. However before the intake operates in a supersonic flow, it must pass through the subsonic flow regime. In some types of supersonic intakes, establishing a shock system with minimal losses is not easy. The process of establishing a stable shock system is referred to as Starting of an intake.

Fig.4.4 Starting of an Intake

External compression intakes complete the supersonic diffusion outside the covered portion of the intake .These intakes usually have one or more oblique shocks followed by a normal shock. Depending upon the location of these shocks, the intake may operate in subcritical, critical or supercritical modes.

4.3 Performance of Intakes

o Performance parameterso Sources of losses o Starting problem in supersonic intakes.o Modes of operation of an external compression intake

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4.4 Boundary Layer Bleed

The formation of a boundary layer on the compression surfaces of the inlet and the interactionOf shocks with the boundary layer cause detrimental pressure gradients and the flow separatesfrom the diffuser walls (Seddon 1988). It will result in an unsteady non-uniform distributionoff low at the engine face, a loss in stagnation pressure recovery, increased internal drag andtoo high Mach numbers due to the reduced flow area at the separation region. This is one ofThe most difficult problems to solve when designing supersonic air-inlet systems

Fig.4.5 Represents Scramjet Flow Path With Boundary Layer Bleed And Corresponding Station Numbering [Ref 5].

Spike bleed air flows through the bleed slot into the translating spike and then enters the front of the fixed centerbody. The air continues through the fixed centerbody, out the four struts, and then flows overboard through louvers. This flow path is illustrated in Figure. [8]

22

Fig.4.6 Spike Bleed Exit Louver[8]

4.5 Bleed System Functions:

4.5.1 Purpose

o Mass flow rate control

o Cabin air quality

o Cooling air systems

o Shock Wave control

Location

o Before Compressor-Supersonic aircraft

Current problems

o Mechanically complex

o Bypass airflow wasted

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Chapter 5

Design and Analysis

To solve any fluid flow or heat transfer problem, one needs to rely on CFD software FLUENT+GAMBIT. It serves various options like conduction, convection, heating, cooling and fluid-structure interaction (FSI) problem. Recently FLUENT is now takeover by ANSYS and named to it as ANSYS CFX. Though people who have already bought FLUENT and GAMBIT separately can use following procedure to analyze a given fluid flow and heat transfer problem via combination of these two software. 

GAMBIT and FLUENT are tools to analysis the fluid flow problems and the branch of science for this problem is known as Computation Fluid Dynamics (CFD) 

5.1 GAMBIT

GAMBIT full form is Geometry And Mesh Building Intelligent Tool

GAMBIT is used for pre-processing operation (which is required before starting of

solution) of fluid flow problem which includes following operations

Geometry creation (specifies the domain of fluid flow problem). It can even model 1D,

2D and 3D domain Mesh generation (discretization of domain to solve governing equations at each cell)

which allows solving every governing equation at each node created in mesh.

Specifying the boundary zones (name & type) to apply boundary conditions for

problem. One need to be very careful in choosing boundary conditions as this will decide

the nature and behavior of any physical phenomena associated with the problem at hand.

GAMBIT export the file containing all the data related to pre-processing.GAMBIT is a type of modeling software and out of it will serve as input to the FLUENT.

24

Fig.5.1 Gambit Topology

5.2 Create Geometry in GAMBIT5.2.1 Create Vertices

The coordinates needed for the mesh are shown below

VERTEX X Y ZA 0 0 0B 2.40 0.756 0C 3.50 1.311 0D 3.70 1.351 0E 3.80 1.367 0F 3.90 1.390 0G 7.00 1.390 0H 8.00 1.362 0I 8.75 1.370 0J 9.12 1.348 0K 9.50 1.320 0L 10.14 0 0M 10.14 1.474 0N 3.50 1.974 0O 10.14 1.974 0P 10.14 2.421 0

Table 5.1: Position of vertex points

25

Fig.5.2 Create real vertex

Using bottom up approach, we start by creating vertices of the geometry using the coordinate given.

Operation Toolpad > Geometry Command Button > Vertex Command Button >

Create Vertex

Create the vertices by entering the coordinates under Global and the label under Label:

26

Click the FIT TO WINDOW button to scale the display so that you can see all the vertices. The resulting image should look like this:

Fig.5.3 Vertex points

5.2.2 Create Edges

Now we can create the edges using the vertices created.

Operation Toolpad > Geometry Command Button > Edge Command Button > Create Edge

Create the edge AB by selecting the vertex A followed by vertex B. Enter AB for Label. Click Apply. GAMBIT will create the edge.

Similarly, create the edges BC, CD, DE, EF, FA , etc., Click on the to select the vertices from the list and move them to the picked list. You can also hold the shift button and mouse click the vertices for selection. The resulting image should look like this.

27

Fig.5.4 Creating Edges

5.2.3 Create Faces

The edges we have created can be joined together to form faces. We will need to define seven faces.

Operation Toolpad > Geometry Command Button > Face Command Button > Form Face

This brings up the Create Face From Wireframe menu. Recall that we had selected vertices in order to create edges. Similarly, we will select edges in order to form a face.

To create the face1, select the edges AB, BC, CE, EF and FA. Enter face1 for the label and click Apply. GAMBIT will tell you that it has "Created face: face1'' in the transcript window.

Similarly, create the face face2 by selecting GH, HI, and GI.

28

Fig.5.5 Creating Faces

5.3 Mesh Geometry in GAMBIT

Operation Toolpad > Mesh Command Button > Edge Command Button ** *>

Mesh Edges*

Select the edge AB. The edge will change color and an arrow will appear on the edge. This indicates that you are ready to mesh this edge. FIRST LENGTH, Select interval size under Spacing. Enter 0.01 for interval

29

Fig.5.6 Mesh edges

Do the same for edge CD and CF.

30

Fig.5.7 Edge mesh

Fig.5.8 Complete edge mesh

31

Now that the appropriate edge meshes have been specified, mesh the face face1:

Operation Toolpad > Mesh Command Button > Face Command Button > Mesh Faces

Select the face1. The face will change color. You can use the defaults of Tri (i.e. triangles) and Pave. Click Apply.

Fig.5.9 Mesh faces

The meshed face should look as follows:

32

Fig.5.10 Face mesh

Next for meshing face2:

Operation Toolpad > Mesh Command Button > Edge Command Button> mesh nodes

Operation Toolpad > Mesh Command Button > Edge Command Button>Link command button

First select edge CF and then select DE click

Apply Second select edge EF and then select CD

click Apply

5.4 Specify Boundary Types in GAMBIT

We'll label the boundary JKLM as pressure far-field, MN pressure inlet, JI,FE as pressure oulet, ABCDE, GHIF as wall and NA as symmetry. Recall that these will be the names that show up under boundary zones when the mesh is read into FLUENT.

5.4.1 Define Boundary Types

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Operation Toolpad > Zones Command Button > Specify Boundary Types

Fig.5.11 Specify boundary types

Save Your Work

Main Menu > File > Save

Export Mesh

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Fig.5.12 Export mesh file

Main Menu > File > Export > Mesh...

Save the file as wedge.msh.

Make sure that the Export 2d Mesh option is selected.

5.5 FLUENT:

This software solve/iterate the problem by importing the file which was exported by

GAMBIT to define the problem in FLUENT. It contains various options to cover all sorts of

problems starting from simple fluid flow problem to complex fluid structural interaction

(FSI).

Fluent will require following data to setup the solution for problem.

Solution method/model

Material properties like density, viscosity, thermal conductivity, specific heat or

internal energy etc. Boundary and operating conditions for given problem so that given physics can be

revealed.

Initial conditions and no. of iterations required to converge the solution. This serves as initial step of algorithm

After applying/specifying above data, FLUENT is ready for solution process. It will

start iterating till the solution gets converged

In FLUENT, solution is performed by use of various solvers which serves as base of

any solution. FLUENT contains various solvers SIMPLE, SIMPLEC, k-epsilon

models or turbulence model etc. 

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After solution is performed; post processing is done for reviewing the results of

solutions to analyze the given problem.

Problem Identification and Pre-Processing 1. Define your modeling goals.

2. Identify the domain you will model.

3. Design and create the grid.

Solver Execution 4. Set up the numerical model. 5. Compute and monitor the solution.

Post-Processing 6. Examine the results. 7. Consider revisions to the model.

Fig.5.13 Gambit- Fluent

5.6 Initial Settings

(Double Click)Setup in the Workbench Project Page.

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When the FLUENT Launcher appears change options to "Double Precision", and then click OK as shown below.The Double Precision option is used to select the double-precision solver. In the double-precision solver, each floating point number is represented using 64 bits in contrast to the single-precision solver which uses 32 bits. The extra bits increase not only

the precision, but also the range of magnitudes that can be represented. The downside of using double precision is that it requires more memory. 

Fig.5.14 Fluent launcher

5.6.1 Problem Setup - General

Now, FLUENT should open. We will begin setting up some options for the solver. In the left hand window (in what I will call the Outline window), under Problem Setup, select General. The only option we need to change here is the type of solver. In the Solver window, select Density-Based.

Fig 5.15 General

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5.6.2 Models

In the outline window, click Models. We will need to utilize the energy equation in order to solve this simulation. Under Models highlight Energy - Off and click Edit.... Now, the Energy window will launch. Check the box next to Energy Equation and hit OK. Doing this turns on the energy equation.

We also need to change the type of viscosity model. Select Viscous - Laminar and press OK

Fig.5.16 Models

5.6.3 Materials

In the Outline window, highlight Materials. In the Materials window, highlight Fluid, and click Create/Edit.... this will launch theCreate/Edit Materials window; here we can specify the properties of the fluid. Set the Density to Ideal Gas, the Specific Heat to1006.43, the Molecular Weight to 28.966. When you have updated these fields, press Change/Create.

Fig.5.17 Materials

5.7 Boundary Conditions

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In the Outline window, select Boundary Conditions. We will now specify each boundary condition for the simulation.

5.7.1 Pressure far-field

In the Boundary Conditions window, select Pressure far-field. Use the drop-down menu to change the Type to pressure-far-field. You will be asked to confirm the change, and do so by pressing OK. Next, a dialogue box will open with some parameters we need to specify. Change the Gauge Pressure (Pascal) to 101325, and Mach Number to 2.95.

Fig.5.18 Pressure far-field

Also, select the Thermal tab, and ensure that the temperature correctly defaulted to 300 K. When you are finished, press OK.

5.8 Operating Conditions

In the Boundary Conditions window, select the Operating Conditions button. Change the Gauge Pressure to 0. Then press OK

Fig.5.19 operating conditions

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It is important to check the operating conditions. When setting the density in materials to ideal gas, FLUENT calculates the density using the absolute pressure. However, the pressure we specify is the gauge pressure, not the absolute pressure. FLUENT will use the absolute pressure to compute the density therefore if we do not set the operating pressure to 0 our density will be incorrect for the flow field.

 5.8.1 Reference Values

In the Outline window, select Reference Values. Change the Compute From parameter to Pressure far-field. Check that the values are accurate. The reference values are used when calculating the non-dimensional results such as the drag coefficient.

Fig.5.20 Reference values

5.9 Solution Methods

In the Outline window, select Solution Methods  to open the Solution Methods window. Under Spatial Discretization, ensure that the option under Flow First Order Upwind is selected.

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Fig.5.21 Solution methods

5.10 Solution Controls

In the Outline window, select Solution Controls to open the Solution Controls window. Ensure that the Courant Number is set to 0.01.

The Courant number can be considered a non dimensionalized timestep. The density-based solver obtains the steady-state solution by starting with the initial guess and marching in pseudo-time until convergence is obtained. The Courant number controls the time step the solver uses. The larger it is, the faster the solution will converge but it will not be very stable and can diverge. The smaller it is, the slower it is to reach convergence but the solution is much more stable.

Fig.5.22 Solution controls

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5.11 Monitors

In the Outline window, click Monitors to open the Monitors window. In the Monitors window, select Residuals - Print,Plot and press Edit.... This will open the Residual Monitors window. We want to change the convergence criteria for our solution. Under Equationand to the right of Continuity, change the Absolute Criteria to 0.001. Repeat for x-velocity, y-velocity, and energy, then press OK.

Fig.5.23 Residual monitors

5.12 Solution Initialization

In the Outline window, select Solution Initialization. We need to make an "Initial Guess" to the solution so FLUENT can iterate to find the final solution. In the Solution Initialization window, select Standard Initialization, then under Compute from, select Pressure far-field from the drop down box. Check to see that the values that generate match our inputted values, and then press Initialize.

Fig.5.24 Solution initialization

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Run Calculation

In the Outline window, select Run Calculation. Change the Number of Iterations to 20000. Double click Calculate to run the calculation. It should a few minutes to solve. After the calculation is complete, save the project. Do not close FLUENT. 

Fig.5.25 Run calculation

RESULTS AND COMPARISON:

Fig.5.26 Scaled Residuals with Slot

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Fig.5.27 Scaled Residuals without Slot

Fig.5.28 (a) Total Temperature Contour with slot

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Fig.5.28 (b) Total Temperature Contour without slot

Fig.5.29 (a) Static Temperature Contour with slot

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Fig.5.29 (b) Static Temperature Contour without slot

Fig.5.30 (a) Static Pressure Contour with slot

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Fig.5.30 (b) Static Pressure Contour without slot

Fig.5.31 (a) Velocity Contour with slot

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Fig.5.31 (b) Velocity Contour without slot

Fig.5.32 Enthalpy Contour with slot

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Fig.5.33 (a) Density Contour with slot

Fig.5.33 (b) Density Contour without slot

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Fig.5.34 (a) Plot of Total Pressure with Slot

Fig.5.34 (b) Plot of Total Pressure without Slot

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Fig.5.35 (a) Plot of Static Pressure with Slot

Fig.5.35 (b) Plot of Static Pressure without Slot

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Fig.5.36 (a) Plot of Static Temperature with Slot

Fig.5.36 (b) Plot of Static Temperature without Slot

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Fig.5.37 (a) Plot of Total Temperature with Slot

Fig.5.37 (b) Plot of Total Temperature without Slot

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Fig.5.38 (a) Plot of velocity with Slot

Fig.5.38 (b) Plot of velocity without Slot

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Fig.5.39 (a) Plot of Enthalpy with Slot

Fig.5.39 (b) Plot of Enthalpy without Slot

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Fig.5.40 (a) Plot of Density with Slot

Fig.5.40 (b) Plot of Density without Slot

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CHAPTER6

Conclusion

Application of a theory which is a solution of the momentum and continuity

equations for inlets with zero angle cowls estimated quite closely the

experimental pressure recoveries obtained. The use of throat bleed in the

vicinity of the center body shoulder controlled flow separation sufficiently to

allow attachment of the internal lip shock. By use of this bleed slot the

performance of the supersonic inlet is varied. The comparison of supersonic

inlet without slot with the supersonic inlet with slot data indicates that with

proper boundary-layer control no severe shock - boundary-layer interaction

losses were incurred by the use of the extreme cowl-lip angles at this Mach

number.

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Chapter 7

References

1. J.Thomas Anderson Details of the Geometry and Operation of the SR-71 Mixed Compression Inlet By Technical Fellow Emeritus Lockheed Martin Skunk Works.

2. Images of F35.

3. D.Kliche ch.Mundt E H Hrischel. The hypersonic Mach number independence principle in the case of viscous flow

4. MATTINGLYElement of Propulsion- Gas turbine and rockets

5. YUE LianJie, XU XianKun & CHANG XinYu. Theoretical analysis of effects of boundary layer bleed on scramjet thrust SDF Aerospace and Aerodynamics Corner

6. Hill and PetersonMechanics and thermodynamics of propulsion

7. Simon trapier Philippe duveau Sebastian deck Experimental study of supersonic inlet buzz.

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