24418180-Study-of-Air-Intake-in-aircraft-report(2).pdf

[Study of air intake in aircrafts] [2009]

[MVJCE, Department of aero] Page 1

A TECHNICAL SEMINAR REPORT ON

“Study of air intake configuration in aircraft”

Submitted in partial fulfillment of requirements for the 1st

semester

MASTER OF TECHNOLOGY

IN

AERONAUTICAL ENGINEERING

Submitted by

CHIRAG.D.SONI

M.Tech, 1st Semester

Dept of Aeronautical Engineering

MVJ College of Engineering

Bangalore-560067



DEPARTMENT OF AERONAUTICAL ENGINEERING

This is to certify that MR. CHIRAG.D.SONI has satisfactorily

completed the seminar of 1st semester Master of technology in

aeronautical engineering prescribed by VTU, Belgaum during the

academic year 2009-2010. The seminar has been approved and

satisfies the academic requirements in respect to the work

prescribed for 1st semester Master of technology.

Name of examiners Signature of HOD

1. ………………………. ………………………..

2. ……………………….



Abstract

Topic: Study of air intake configurations in aircraft.

The air intake is that part of an aircraft structure by means of

which the aircraft engine is supplied with air taken from the

outside atmosphere. The air flow enters the intake and is

required to reach the engine face with optimum levels of total

pressure and flow uniformity. These properties are vital to the

performance and stability of engine operation. Depending on the

type of installation, this stream of air may pass over the aircraft

body before entering the intake properly.

Selection of the correct type of intake and the associated

inlet geometry has important consequences to any airplane

design. For that reason, intake design receives considerable

attention in the design phase of an airplane.



Contents

Chapter1 Page no:

1.1 Introduction to air intake 1

1.2 Need of air intake system 2

1.3 Air intake Design requirements 3

1.4 Intake configurations 4

Chapter 2

2.1 Jet engine intake (subsonic) 11

2.2 Determination of size of the stream tube 15

2.3 Deceleration of airflow 16

2.4 Air intake characteristics of Lockheed C-141 19

Chapter 3

3.1 Jet Engine Intakes: Supersonic 22

3.2 Flow conditions over wedge and cone 26

3.3 Intake configuration and operation 30

3.4 Examples of oblique shock diffusers 34

3.5 Supersonic air intake case studies 36

References 41



List of figures

Chapter 1 page no:

1.1 Air intake in aircraft 2

1.2 Turboprop engine air intake 4

1.3 Plenum Inlet 5

1.4 Subsonic Bifurcated Inlet 6

1.5 Subsonic Podded Nacelle Inlet 7

1.6 Pitot type intake 8

1.7NACA Submerged Inlet in a Euro Fighter 9

Chapter 2

3.1 Intake flow field 12

3.2 Intake flow field at high speed 17

3.3 Air intake in Lockheed C-141 19

Chapter 3

3.10 Supersonic flow over wedge and cone 26

3.12 Comparison of supersonic flow 28

Over wedge and cone



3.11Total pressure loss and static pressure 29

Increase due to shockwave.

3.13 Operation of normal shock diffuser 31

3.15 Characteristics of oblique shock diffuser 33

3.16 Examples of oblique shock diffusers 34

3.17 F-16 intake characteristics 38

3.25 F-14 intake characteristics 41



Introduction



Chapter 1

INTRODUCTION

1.1What is air intake?

In any application subsonic transport or supersonic fighter the air intake is

essentially a fluid flow duct whose task is to process the airflow in a way that

ensures the engines functions properly to generate thrust.

Fig 1.1[air intake system]

1.2 Need of air intake in an aircraft.

A widely used method to increase the thrust generated by the aircraft

engine is to increase the air flow rate in the air intake by using auxiliary

air intake systems.

The air flow enters the intake and is required to reach the engine face

with optimum levels of total pressure and flow uniformity hence need of

an air intake system.

Deceleration of airflow at high flight mach numbers or aerodynamic

compression with help of air intake.



1.3 Air intake design requirements

The airflow first passes through the air intake when approaching the engine,

therefore the intake must be designed to meet certain requirements of aircraft

engines such as:

The air intake requires enormous effort properly to control airflow to the

engine.

The intake must be designed to provide the appropriate amount of

airflow required by the engine.

Furthermore this flow when leaving the intake section to enter the

compressor should be uniform stable and of high quality.

Good air intake design is therefore a prerequisite if installed engine

performance is to come close to performance figures obtained at the

static test bench.

The engine intake must be a low drag, light weight construction that is

carefully and exactly manufactured.

These above conditions must be met not only during all phases of flight

but also on the ground with the aircraft at rest and the engine demand

maximum, thrust prior to take off



1.4 INTAKE CONFIGURATIONS

Broadly the intake configurations may be classified as

1. Piston engine intakes

2. Turbo propeller intakes

3. Jet Engine Intakes: Subsonic

4. Jet Engine Intakes: Supersonic

Jet Engine Intakes: Subsonic

These are of the following types:

1. Plenum Intake

2. Bifurcated Intake

3. Podded nacelle Inlet

4. Pitot Inlet

5. NACA Submerged Inlet

Turboprop engine air intake as seen below fig [1.2]



Subsonic intakes

Plenum Intakes

These are used mainly in combination with double-sided centrifugal flow

compressors. In this case the engine is installed in a region of large volume, the

‘plenum chamber’, in order that front and rear compressor intakes can receive

equal air supplies. The aircraft intake feeds directly into the plenum chamber.

Fig 1.3 shows a sectional view of plenum intake.

Fig 1.3 Plenum Inlet



Bifurcated intakes are used primarily in single engine installations with side

intakes Fig 1.4 shows a bifurcated intake.

Fig. 1.4 Subsonic Bifurcated Inlet



Fig 1.5 Subsonic Podded Nacelle Inlet

.



Pitot type intakes have been applied to many fighter airplanes. They are not

influenced by the flow field of other airplane components. However, they

require very long ducts which cause extra weight and loss in pressure recovery.

Fig 1.6 shows a pitot type intake



The NACA submerged type intake is not very efficient for use with

propulsion installations. However, they are frequently used as intakes of

auxiliary systems (auxiliary power unit, heating and avionics bay cooling) as

seen in Fig 1.7

Fig 1.7NACA Submerged Inlet in a Euro Fighter

Except for the Pitot and the Podded nacelle type intakes, all jet

engines intakes are equipped with boundary layer diverters (or B.L. Splitters).

If such boundary layer diverters are not used, large pressure recovery losses

(thus losses in thrust) are incurred.



A major consideration in jet fighter intake design is the behavior of the intake

at very high angles of attack and sideslip. Compressor stall and engine surging

are easily induced in such conditions.

In subsonic installations, the intake is kept as short as possible. Long

ducts translate into weight and pressure recovery losses. In jet fighters and in

jet trainers long ducts cannot always be avoided.



Jet engine intake

(subsonic)



2.1Subsonic air intakes

The standard air intake has found widespread application with high subsonic

civil and transport aircraft. Being of quasi circular cross section, the air intake

forms the forward part of the engine nacelle. Subsonic air intakes are also

applied to some combat aircrafts and virtually all jet training aircrafts that

operate near the speed of sound. Here we find intake shapes of elliptical ,half

circular ,or even irregular cross section ,with intake mounted on the fuselage

sided or under the fuselage .





Observed, the bounding streamlines of which will terminate in stagnation

points on the cowl. With aircraft velocity increasing, stagnation points continue

to move forward of the cowl.

2.2 Determination of size of stream tube

Cross section A0, of the stream tube well ahead of the intake is determined by

the engine mass flow rate, the size of the stream tube may simply be

determined by applying the continuity considerations. Continuity requires mass

flow rate m. at any cross section within the stream tube to be the same, which is

hence a constant. Mass flow rate at cross-section A0, in particular ,exactly

equals mass flow rate at the compressor face A=2=,which itself reflects engine

mass flow .hence:

m.0=m

.2

Mass conservation may be expressed for the a particular flow path station

(upstream infinity) and 2(compressor face) as follows

Station 0(upstream infinity)

m.0=p0v0A0

Station 2 (compressor face):

m.2=p2v2A2

Therefore cross section of the stream tube at upstream infinity will result as

simple expression:

A0= (p2/p0)*(v2/v0)*A2

If air density is assumed not to change within the stream tube between the

stations 0 and 2 ,then stream tube cross-section A0 depends only on aircraft

flight speed v0 , because air stream velocity at compressor face is determined



by the compressor ,with compressor entrance cross-section A2 a constant by

design.

2.3 Deceleration of airflow at high flight mach numbers or

aerodynamic compression with help of air intake.

We know that for an air breathing engine to function correctly compression of

air is a prerequisite. Aerodynamic compression occurs in flow ducts whose

cross-sectional area gradually increases in stream wise direction. A duct with

the ability to retard the flow and convert energy into pressure energy is termed

as diffuser.

At sufficiently high mach numbers, for instance at cruising flight, airflow

approaching the engine will be faster then would be tolerable for the

compressor. Due to the diffuser action of air intake which is deceleration of the

air flow and a buildup of pressure, airstream velocity will be adapted to the

need of the compressor as seen in fig 3.2a. Additionally, due to the rise in

pressure, a considerable benefit to the engine cycles results so that less

mechanical energy is required for compression.

Pressure recovery and nose suction

In order to prevent the flow from separating along the walls , the interior

surface of the diffuser must be carefully shaped , and be smooth and

unobstructed by steps or kinks , otherwise the sensitive boundary layer

(between main stream and diffuser wall ) may separate. This would result in

partial losss of kinetic energy and its conversion into unusable heat, a process

termed friction which always results in a degradation of total pressure. If it

were possible for the deceleration flow to convert all its of kinetic energy into

pressure , then total pressure of the flow would remain constant and so-called

pressure recovery would be 100%



Nose suction formation

Today’s high-subsonic cruise flight mach numbers which are in range of mach

0.78 to 0.85 call for an air intake design which features a relatively ‘thin’ intake

i.e. where external dimension of intake is not much greater than the internal

diameter. This will result in a small radius, leading to a relatively thin lipped air

intake.

If the external flow is made to pass the intake lip ‘correctly’, additional drag

resulting from ram effect ahead of the intake may effectively be reduced. Such

a reduction is accomplished solely by the air stream flowing around the nose.

As the flow follows the contour of the nose, excessive velocities can develop

which may even attains (low) supersonic speeds. This will cause a zone of low

pressure around the intake‘s circumference , leading to the exertion of an

aerodynamic force with a component acting in the direction of engine thrust

and termed as nose suction [3.2b].





2.4 Air intake characteristics of Lockheed C-141 strlifter

military transport

The intake is particularly noteworthy because of its short duct, denotes

as ‘zero-length inlet’ by Lockheed, which enabled a light weight

constructions of high aerodynamic performance (fig 3.3).

Due to its small radius, the intake lip is relatively sharp-edged which

made necessary a secondary intake system that comes into effect at high

airflow rates with aircraft static , or at low speed.

The slotted inlet embodies 12 sets of outer doors pivoted at the cowl.

The door opens against a spring force if a [pressure drop exists between

the low static pressure on the engine side of doors relative to that of the

external side of the doors.







Jet engine intake

(supersonic)



Chapter 3

3.1 Jet Engine Intakes: Supersonic

They are of the following types:

1. Pitot Intake

2. External compression Intake

3. Mixed (or external/internal) compression Intake

A Pitot Intake has a number of attractive features, notably low drag and

a stable flow characteristic with good flow distribution. Its disadvantage lies in

the level of pressure recovery achieved. As shown in Fig 1.6, this type of intake

has been used in aircrafts like the Mig 21.

Fig 1.6 Mig-21 Air Intake



Proper inlet design is extremely critical to supersonic aircrafts. A long

inlet duct is often needed to assure smooth flow deceleration (to around M=0.4

at the compressor face) and to assure full use of the favorable pressure

distribution in the inlet duct. A typical intake for a twin engine aircraft is

shown Fig 1.7. Different types of supersonic intakes are given in Fig 1.8 and

some examples of supersonic intakes are shown in Fig 1.9.

Fig 1.7 Supersonic Twin Engine Inlet



Fig 1.8 Supersonic Inlets



Pressure waves in air



3.2 Flow conditions over wedge and cone

In the design of supersonic air intakes flow conditions over wedge and cone are

of the greatest importance as these are simple geometric bodies and relatively

easy to manufacture.

First let us consider supersonic flow over a wedge. Such a device is installed in

the air intake of the majority of modern supersonic combat aircraft such as F-

15 F-14, MiG-29, Su-27, but also in the airliner Concorde.

We assume a wedge of unlimited length to be latterly immersed in a

supersonic gas stream (fig 3-10a). Flow conditions here are similar to the



previously discussed corner flow where streamlines, after passing the shock

front, are everywhere tangent to the wedge cross-section. Due to the

compressive effect of the shock, the stream line pattern downstream of the

shock is more compact hen it is upstream.

If the wedge angle exceeds the maximum value permissible for that particular

Mach number, the oblique shock will no longer remain attached but will jump

abruptly upstream to form a (detached) bow shock. Part of the bow shock

immediately ahead of the wedge apex acts like a normal shock causing the

region between shock and wedge to be sub sonic, i.e. M<1 (fig 3-10b) adjacent

regions of the shock surface bounding the center normal shock region,

increasingly bend in a downstream direction to form an oblique shock with,

finally, degenerates into a (weak) mach line (not shown) .

In order to design aircraft of low wave drag, the angle of the shock front must

be small. This implies, apart from the supersonic Mach number flown, that

nose sections of intake and wing must be given a knife-edge shape. We now

understand why subsonic intakes with their well rounded nose sections are of

less use in supersonic flow: the detached bow shock creates high drag which

will absorb much of the engine’s thrust, so that supersonic flight is virtually

unattainable.

Comparison of supersonic flow over cone and wedge

The major advantage of a (supersonic) conical flow is a smaller total pressure

loss (when compared to a wedge of the same half-angle), together with the fact

that a conical shock sustains lower mach numbers until it becomes detached to

form a high loss bow shock.

A major disadvantage of conical flows is that it is less tolerant to asymmetric

flow conditions which cause distortion to the intake flow. As combat aircraft

are frequently required to maneuver at higher angles of attack, the flow

inevitably gets asymmetric- hence a performance for the (horizontally

arranged) wedge in all modern combat aircraft, despite its reduced efficiency.





Total pressure loss and static pressure increase due to

shockwave.



3.3 Intake configuration and operation

Present-day turbine aero engines require subsonic flow at the entry to the

compressor, even if the aircraft is flying at supersonic speed. The task of air

intake is therefore to decelerate the supersonic external flow to a subsonic

speed acceptable to the compressor. As intake discharge mach number are

required to be in range of mach 0.4 to 0.7 great care must be exercised when

decelerating the flow in order to keep total pressure losses to a minimum .

Normal shock diffuser

For aircraft operating at a maximum speed equivalent to mach 1.5 a normal

shock diffuser is generally sufficient to decelerate the supersonic airflow

efficiently to the speed needed by the compressor.

The action of diffusing i.e. the deceleration of flow and build up of pressure is

accomplished in two steps:

The supersonic flow is (abruptly) decelerated, through the normal shock

, to subsonic velocity with an accompanying abrupt increase in static

pressure;

In the diverging (subsonic) duct, where the flow is sill faster then would

be acceptable to the compressor, deceleration of the flow continues with

pressure increasing further.

Case 2

Suppose the air flow demand of the engine is reduced. Then static pressure p2

at the compressor face will rise ,less air is allowed to enter the intake, the

excess airflow after being processed through the shock front is forced to flow

outside the inlet as a so-called spill-over flow (fig 3-13b).

Case 3

Suppose the air flow demand of the engine to be greater than the intake can so

provide. At first, this is equivalent o pressure drop at the compressor inlet,

either pressure decreasing upstream, too. This will eventually cause the shock



to be swallowed, and the airstream to enter the subsonic diffuser at supersonic

velocity .The inconsistency of duct geometry and flow velocity results in a

complex shockwave pattern within the duct (fig 3.13c).



Oblique shock diffuser intake characteristics

The operational characteristics of an oblique shock diffuser may be

summarized in three typical conditions.

Case1: If the normal shock that terminates the supersonic flow regime is

exactly at the position of the diffuser throat (i.e. where the cross section is a

minimum), the airflow rate is a maximum fig (3.15). This condition is denoted

as critical. The inclination angle of the first oblique shock wave is then

determined both by the free stream mach no and the apex angle of wedge or

cone. Such a shock configuration assures acceptable intake efficiency and

usually corresponds to the design pint of the diffuser.

Case 2: In case of a pressure drop at the compressor face, the normal shock

will be swallowed to adopt a quasi-stable position farther down-stream within

the intake duct (fig 3-15c). This condition is denoted as supercritical and, due

to greater strength of the (terminating) normal shock, poor flow quality results.

Case 3: Now assume a rise in the pressure at the compressor face such as

caused by a reduced airflow demand of the engine. The normal shock will then

be expelled from its throat position, air flow is reduced. Intake operation in this

case is subcritical (fig 3-15b).such a shock position is highly unstable, the

shock oscillating at high rate between swallowed and expelled positions. This

oscillating motion causes high frequency pressure oscillations in the intake;

known as diffuser buzz a sound feared by pilot’s as it can indicate one of the

most dangerous conditions of the propulsion system.





3.4 Examples of oblique shock diffusers

Mirage ||| fighter with side mounted oblique-shock diffuser

fig (3.16a)

Axisymmetric oblique-shock diffuser (Lockheed SR-71) fig

(3.16b)



Two dimensional oblique shock diffuser (Northrop F-5 with

vertical ramp) Fig (3.16c)



3.5 Supersonic air intake case studies

An aircraft showing the typical application of a normal shock diffuser is the

American F-16, now a product of Lockheed, but developed and built originally

by the general dynamics corporation.



The F-16 intake is of fixed-geometry type, without moveable parts a

decision made fairly in design process to save costs.

What is remarkable about the inlet is its positioning fairly well under the

fuselage a solution resulting from the requirements of aircraft

The F-16 was designed to have exceptional maneuverability and this

required to operate at high angle of attack. In these considerations the

long fuselage fore body performs a shielding function which serves to

align the (inclined) axis of intake (fig 3-17a).

The intake itself features a short duct which not only contributes to the

light weight design of the aircraft, but also minimizes flow distortion

ahead of the compressor.

Another problem facing the combat aircraft is the hot gas from gun

muzzles that may be ingested and cause engine flame out. By placing

the gun muzzle above the leading-edge extension or strake, the high

temperature gas from the gun will be kept effectively away from the

intake before being carried away by the external flow as shown in fig 3-

17 below.

The intake cowl features a moderately blunt lower lip that transitions

into a sharp leading-edge extension or splitter plate on the upper side

(close to the fuselage). The splitter plate extends 25cmsahead of the

lower cowl lip to isolate the inlet normal shock from the fuselage

boundary layer (fig-17b).

A short length of splitter plate keeps boundary layer buildup small, so

eliminating the need of boundary layer bleed on the splitter.



Fig (3.17)



Intake characteristics of F-14





References

Jet engine fundamentals theory and operations by

KLAUS HUNECKE.

Cowl - Wikipedia, the free encyclopedia.

Aircraft engine controls - Wikipedia, the free

encyclopedia.

Turboprop - Wikipedia, the free encyclopedia.

‘Janes’ All the World’s Aircrafts ’, 2000.

http://en.wikipedia.org/wiki/Cowl

http://en.wikipedia.org/wiki/Aircraft_engine_controls

http://en.wikipedia.org/wiki/Aircraft_engine_controls

http://en.wikipedia.org/wiki/Turboprop

24418180-Study-of-Air-Intake-in-aircraft-report(2).pdf

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