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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
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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. ……………………….
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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.
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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
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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
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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
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Introduction
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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.
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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
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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]
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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
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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
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Fig 1.5 Subsonic Podded Nacelle Inlet
.
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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
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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.
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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.
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Jet engine intake
(subsonic)
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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 .
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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
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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%
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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].
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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.
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Jet engine intake
(supersonic)
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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
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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
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Fig 1.8 Supersonic Inlets
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Pressure waves in air
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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
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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.
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Total pressure loss and static pressure increase due to
shockwave.
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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
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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).
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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.
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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)
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Two dimensional oblique shock diffuser (Northrop F-5 with
vertical ramp) Fig (3.16c)
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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.
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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.
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Fig (3.17)
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Intake characteristics of F-14
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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.