Subsonic and supersonic air intakes

SUBSONIC AND

SUPERSONIC AIR INTAKES

(JET PROPULSION)

SANJAY SINGH

Asst. Prof. and Head

Department Of Aeronautical Engineering

VMKV Engineering College

Salem (Tamilnadu)

[email protected]

INTRODUCTION

• Inlets are very important to the overall jetengine performance & will greatlyinfluence jet engine thrust output.

• The faster the airplane goes the morecritical the inlet duct design becomes.

• Engine thrust will be high only if the inletduct supplies the engine with the requiredairflow at the highest possible pressure.

• The nacelle/duct must allow the engine tooperate with minimum stall/surgetendencies & permit wide variation inangle of attack & yaw of the aircraft.

• For subsonic aircraft, the nacelle shouldn’tproduce strong shock waves or flowseparations & should be of minimumweight for both subsonic & supersonicdesigns.

• Inlet ducts add to parasitic drag (skinfriction+ viscous drag) & interference drag.

• It must operate from static ground run upto high aircraft Mach number with highduct efficiency at all altitude, attitudes &flight speeds.

• It should be as straight & smooth aspossible & designed in such a way thatBoundary layer separation is minimum.

• It should deliver pressure distributionevenly to the compressor.

• Spring loaded , Blow-in or Suck-in- Doors aresometimes placed around the side of theinlet to provide enough air to the engine athigh engine rpm & low aircraft speed. It isalso operated during compressor surge /stall.

• It must be shaped in such a way that ramvelocity is slowly & smoothly decreases whilethe ram pressure is slowly & smoothlyincreases.

SUBSONIC INLETS

Types

• 1. Internal Compression Subsonic Intakes

• 2. External Compression Subsonic Intakes

Internal Compression Subsonic Intakes

• A divergent duct acts as a subsonic internalcompression diffuser.

• The pressure gradients of this intakes are keptlow enough to avoid large stagnation pressureloss.

• To keep a low pressure gradient, thedivergence angle must be made small whichincreases the length of the diffuser and hencethe associated friction loss.

Internal Compression Subsonic Intakes

External Compression Subsonic Intakes

• As we know that boundary layer in the diffuser passage leads to losses, if the compression of the gas is made to occur before it enters the diffuser passage (i.e. external to the diffuser), near isentropic compression is possible.

• The inlet is made up of a constant area duct enclosed by a contoured cowl.


• The presence of the cowl causes the stagnation stream lines to diverge between the upstream and the inlet causing compression between the two sections.

• Such inlets are not suitable for high subsonic Mach No. applications due to the possibility of local Mach No. greater than 1.


SUBSONIC DUCTS

BOUNDARY LAYER

Inlet design

• Inlet design requires a compromise betweenexternal and internal deceleration.

• Both can lead to difficulties, and a balance isneeded.

• To examine the effect of externaldeceleration on inlet design, methods areneeded for calculating both potential flow(internal and external) and boundary layergrowth on intake surfaces.

Boundary Layer Separation• In an actual engine inlet separation can take

place in any of the three zones as shown in figure.

• 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.

Boundary Layer Separation• 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, then decelerates as the curvature decreases.

Boundary Layer Separation• In some installations, it has not been possible to

make the exit area of the intake more than about 30% greater than the inlet area without the incidence of stall and large losses.

• Reynolds number effects is also important for large inlets and high – speed flow.

• At high angles of attack, all three zones are subjected to unusual pressure gradients.

Thrust on inlet Surface

Thrust on inlet Surface

• Net momentum flux out of the control volume is

• Net momentum flux is

• Bernoulli’s law

The above relation shows that the

greater external deceleration (i.e.

the smaller the ratio ui / ua ), the

larger must be thrust increment.

Coefficient of Pressure (Cp)

• On the outer surface of the nacelle, thepressure must rise from some minimumvalue Pmin (at the point where the local free– stream velocity is umax ) to the ambientvalue Pa associated with straight parallelflow downstream neglecting boundarylayer.

Coefficient of Pressure (Cp)• Cp must not be too large otherwise the

boundary layer will separate.

• An average pressure difference with a factor ‘s’value of which lies between 0 to 1 can bewritten as

• Therefore, thrust increment equation can be written as

Area Ratio

• Therefore the area ratio can be expressed in terms of external deceleration ratio.

External Deceleration

• From the relation of Area Ratio and ExternalDeceleration, it is clear that the larger theexternal deceleration (the smaller the value ofof ui / ua), the larger must be the size of thenacelle, if one is to prevent excessive drag.

• Even in the absence of boundary layerseparation, the larger the nacelle, the larger theaerodynamic drag on it.

• If the external deceleration is modest ( e.g.ui / ua > 0.8), its effect on minimum nacellesize is quite small.

Internal Deceleration

• The use of partial internal deceleration ismore effective in reducing maximumdiameter because it permits a reduction inboth Ai and Amax / Ai .

• Performance of an inlet depends on thepressure gradient on both internal andexternal surfaces.

Pressure Rise (External & Internal)

• External pressure rise is fixed by the externalcompression and the ratio Amax / Ai .

• Internal pressure rise depends on thereduction of velocity between entry to the inletdiffuser and entry to the compressor (or burnerfor a ramjet).

• Nacelle size required for low drag can be quitestrongly dependent on the degree of externaldeceleration.

Inlet Performance Criterion

Performance Criterion

1. Isentropic Efficiency of a Diffuser (defined in terms of temperature rise).

State 02s is defined as the state that would bereached by isentropic compression to the actualoutlet stagnation pressure.

• Since,

• Diffuser efficiency can be written as

Ram Efficiency

2. Ram Efficiency (Defined in terms of pressure rises)

ᶯr = (P02 - Pa ) / P0a - Pa

Stagnation Pressure ratio• The Stagnation Pressure ratio , rd is widely

used as a measure of diffuser performance.

• For supersonic intakes

rd = 1 – 0.75 (Ma - 1)1.35

(A rough working rule adopted by American Dept. of Defense)

Which is valid when 1 < M < 5.

• To obtain the overall pressure recovery

factor, P02 / P0a must be multiplied by the

pressure recovery factor for the subsonic part ofthe intake.

• Diffuser efficiency and Stagnation Pressure ratio are related.

• The relationship between internal and externaldeceleration depends on engine mass flow rateas well as flight Mach number M.

DUCT EFFICIENCY

• The duct pressure efficiency ratio isdefined as the ability of the duct to convertthe kinetic or dynamic pressure energy atthe inlet of the duct to the static pressureenergy at the inlet of the compressor withouta loss in total pressure .

• It is in order of 98% if there is less frictionloss.

RAM RECOVERY POINT

• The Ram Recovery Point is that aircraftspeed at which the ram pressure rise isequal to the friction pressure losses

OR

• That aircraft speed at which thecompressor inlet total pressure is equalto the outside ambient air pressure.

• A good subsonic duct has aircraft speed of257.4 km/h for a good ram recovery point.

Supersonic Inlets

• Even for supersonic flight it remainsnecessary that the flow leaving the inletsystem be subsonic.

• It is required to have some means todecelerate supersonic flow to subsonicspeeds tolerable by existing compressorsor fans.

Types of Supersonic Inlets

• Reverse Nozzle Diffuser or

Converging - Diverging

Intakes

• Normal Shock Diffuser or Pitot

Inlet

• Oblique Shock Diffuser

Reverse Nozzle Diffuser or Diffusers

with internal contraction or Converging Diverging Intakes

• Deceleration from supersonic to subsonicflow speeds can be done by a simplenormal shock with small stagnationpressure loss if the upstream Machnumber is close to 1.

• For high Mach number the loss across asingle normal shock would be excessive.

• Therefore it is better to use a combinationof oblique shocks.

Normal-Shock diffuser• All existing compressors and fans require

subsonic flow at their inlet with 0.5 < M2 < 0.8at high power conditions.

• So the inlet must reduce the flow Machnumber from Mo > 1 to M2 < 1.

• The simplest way to do this is with a NormalShock.

• Prandtl – Meyer Relation for the normalshock in a perfect gas is

• V 1V 2 = a*2 = 2a02 / ᵞ + 1

• M1* M2

* = 1

Normal-Shock diffuser

Normal-Shock diffuser

• For low supersonic speeds, such diffusersare adequate because the stagnationpressure loss is small, but at Mo = 2, pt2 /pto ≈ 0.71, a serious penalty, and at Mo = 3pt2 / pto ≈ 0.32.

• For example the F-16 fighter has a simplenormal shock diffuser, while the F-15 hasan oblique shock diffuser.

• The losses can be greatly reduced bydecelerating the flow through one or moreoblique shocks, the deflection and the pressurerise of each being small enough to be in therange where the stagnation pressure ratio isclose to unity.

• It is very important to understand that anOblique Shock is in fact just a normal shockstanding at an angle to the flow.

Oblique - Shock diffusers



• M1n is given in terms of Mon by the samerelation given for M1 as a function of Mo. ButMon can be made close to 1.

• The condition for a weak sound wave is justMon = 1,


• By choosing the wedge angle (ordeflection angle) ∂ we can set the shockangle.

• A series of weak oblique shocks, for eachof which the Mn is near unity, hence alllying in the range of small pt loss, can yieldan efficient diffuser.

SUPERSONIC INLETS WITH

VARIABLE GEOMETRY

• This would work at one design Mach number,the one for which the isentropic area ratiobetween the incoming supersonic flow and thesonic throat is exactly the as-built area ratio A1 /Athroat .

• But during the acceleration to this Mach numberthe fully supersonic flow cannot be establishedin the inlet without varying the geometry.

• Imagine the inlet flying at M0 , lower than thedesign Mach number.

• The flow will look as depicted in the top right indiagram shown in next slide.

This is because at the lower M0 the flow areathat would decelerate isentropically to sonic atthe throat is smaller than the built area A1.

STARTING THE DIFFUSER

• If the flow arrives undisturbed at the inlet,it could only occupy a fraction of it, the restof the flow into the frontal area A1 isrequired to be disposed of which is called“Spillage”.

• This “Spillage” is accomplished by thedetached normal shock; behind it the flowis subsonic and it can turn around theinlet.

• The shock at the full flight Mach number is verylossy, and it is not practical to simply force theplane to continue accelerating to the designcondition (there may not even be enough thrustleft to do it).

• What can be done is to manipulate thegeometry to swallow the shock and reduce itsstrength. This is called “STARTING” THEDIFFUSER.

STARTING OF THE DIFFUSER

• To "START" THE DIFFUSER, means to passthe shock through the convergent portion, thereshould be an increase in the throat area untilthe normal shock is just at the lip.

• At that point, any further small increase in throatarea causes the shock to jump rapidly to aposition in the divergent part of the nozzlewhere the area is again A1.

• This rapid jumping of shock from convergingportion to diverging portion takes place becausethe shock is unstable in the converging section,but stable in the divergent section.

• This is accomplished by the flow due to whichrepositioning of the shock to a location nearerthe throat, on the supersonic side takes place.

• The process can continue until the shock isalmost at the throat.

• This repositioning of shock in throat on thesupersonic side is called “STARTING OF THEDIFFUSER” . For this successive steps ofacceleration is followed.

Successive Steps in Acceleration of a CD inlet

Condition (a)

• Low subsonic speed operation.

• Inlet is not choked.

• The airflow through the inlet and hence theupstream capture area Aa is determinedby conditions downstream of the inlet.

Condition (b)• Low subsonic speed operation.

• Inlet is choked.

• The inlet mass flow rate is limited by thechoking condition at At .

• Since the flow is isentropic, At = A* and theupstream capture area Aa + is given by

Condition (c) to (f)

• In condition (c) to (f), the inlet flow velocity isincreased gradually to design Mach No. MD toposition the shock first in front of the inlet, thenin the cowl or inlet lip, then it enters in theconverging part and jumps rapidly to divergingpart.

• When the air intake starts operating in designMach No. MD, the shock repositions itself to thethroat nearer to supersonic side.

MODES OF INLET OPERATION

CRITICAL INLET 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.

Modes of Inlet Operation

SUB - CRITICAL INLET 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.


SUPER - CRITICAL INLET OPERATION

• The condition when the inlet can

not capture the mass flow

required by the engine and the

terminal shock is sucked into the

diffuser is called super - critical

operation.


FLOW INSTABILITYBUZZ

• Buzz is an airflow instability caused by theshock waves rapidly being alternatelyswallowed and expelled at the inlet of the ductand occurs in supersonic intakes at subcriticaloperations.

• It starts when the aircraft begins to fly at or nearthe speed of sound. At these speeds sonicshock waves are developed that if notcontrolled will give high duct loss in pressureand airflow and will set up vibrating conditionsin the inlet duct, called inlet Buzz.

Variable Geometry DuctAt higher Mach No., the inlet duct geometry is made

variable by any one of the following:

• (a) Moving the inlet spike in and out so as to

maintain the oblique shock on the edge of the outer

lip of the duct (axisymmetric duct).

• (b) Moving the side wall or ramp to a higher angle so

as to force a stronger oblique shock front (2-

dimensional duct).

• (c) Varying the normal shock (expanding centre

body).

• (d) Varying the inlet lip area so as to vary the intake

area.

Rating of Engines - Bell Mouth Inlet

• The manufacturers rate their engines using abellmouth inlet. It is a subsonic inlet.

• This type of inlet is essentially a bell shapedfunnel having carefully rounded shoulderswhich offers practically no air resistance.

• The duct loss is so small that it is consideredzero and all engine performance data can begathered without any correction for inlet ductloss being necessary.

THANK YOU.

Technology

Subsonic and supersonic air intakes