Subsonic And Supersonic Air Intakes a322y

SUBSONIC AND SUPERSONIC AIR INTAKES (JET PROPULSION)

INTRODUCTION • Inlets are very important to the overall jet engine performance & will greatly influence jet engine thrust output. The faster the airplane goes the more • critical the inlet duct design becomes. • Engine thrust will be high only if the inlet duct supplies the engine with the required airflow at the highest possible pressure.

• The nacelle/duct must allow the engine to operate with minimum stall/surge tendencies & permit wide variation in angle of attack & yaw of the aircraft.

For subsonic aircraft, the nacelle shouldn’t • produce strong shock waves or flow separations & should be of minimum weight for both subsonic & supersonic designs.

• Inlet ducts add to parasitic drag (skin friction+ viscous drag) & interference drag. It must operate from static ground run up • to high aircraft Mach number with high duct efficiency at all altitude, attitudes & flight speeds. • It should be as straight & smooth as possible & designed in such a way that Boundary layer separation is minimum. • It should deliver pressure distribution evenly to the compressor.

• Spring loaded, Blow-in or Suck-in- Doors are sometimes placed around the side of the inlet to provide enough air to the engine at high engine rpm & low aircraft speed. It is also operated during compressor surge / stall. • It must be shaped in such a way that ram velocity is slowly & smoothly decreases while the ram pressure is slowly & smoothly increases.

SUBSONIC INLETS Types • 1. Internal Compression Subsonic Intakes • 2. External Compression Subsonic Intakes

Internal Compression Subsonic Intakes • A divergent duct acts as a subsonic internal compression diff. The pressure gradients of these intakes are • kept low enough to avoid large stagnation pressure loss. • To keep a low pressure gradient, the divergence angle must be made small which increases the length of the diff and hence the associated friction loss.

Internal Compression Subsonic Intakes

External Compression Subsonic Intakes • As we know that boundary layer in the diff age leads to losses, if the compression of the gas is made to occur before it enters the diff age (i.e. external to the diff), near isentropic compression is possible. • The inlet is made up of a constant area duct enclosed by a contoured cowl.

External Compression Subsonic Intakes • 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.

External Compression Subsonic Intakes

SUBSONIC DUCTS

BOUNDARY LAYER

Inlet design • Inlet design requires a compromise between external and internal deceleration. • Both can lead to difficulties, and a balance is needed. To examine the effect of external deceleration • on inlet design, methods are needed for calculating both potential flow (internal and external) and boundary layer growth 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, and 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. A Reynolds number effect 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 () • On the outer surface of the nacelle, the pressure must rise from some minimum value Pmin (at the point where the local free – stream velocity is umax ) to the ambient value Pa associated with straight parallel flow downstream neglecting boundary layer.

Coefficient of Pressure ()

• 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 be written as • Therefore, thrust increment equation can be written as

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

External Deceleration • From the relation of Area Ratio and External Deceleration, it is clear that the larger the external deceleration (the smaller the value of ui / ua), the larger must be the size of the nacelle, if one is to prevent excessive drag. • Even in the absence of boundary layer separation, the larger the nacelle the larger the aerodynamic drag on it. • If the external deceleration is modest (e.g. u i / ua > 0.8), its effect on minimum nacelle size is quite small.

Internal Deceleration • The use of partial internal deceleration is more effective in reducing maximum diameter because it permits a reduction in both Ai and Amax / Ai . • Performance of an inlet depends on the pressure gradient on both internal and external surfaces.

Pressure Rise (External & Internal) • External pressure rise is fixed by the external compression and the ratio Amax / Ai . • Internal pressure rise depends on the reduction of velocity between entry to the inlet diff and entry to the compressor (or burner for a ramjet). • Nacelle size required for low drag can be quite strongly dependent on the degree of external deceleration.

Inlet Performance Criterion

Performance Criterion 1. Isentropic Efficiency of a Diff (defined in of temperature rise).

State 02s is defined as the state that would be reached by isentropic compression to the actual outlet stagnation pressure.

• Since,

• Diff efficiency

can be written as

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

ᶯr = (P02 - Pa ) / P0a - Pa

Stagnation Pressure ratio • The Stagnation Pressure ratio , rd is widely used as a measure of diff 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,

P

P

/ must be multiplied by 0a the pressure recovery factor for the subsonic part of the intake. 02

• Diff efficiency and Stagnation Pressure ratio are related.

• The relationship between internal and external deceleration depends on engine mass flow rate as well as flight Mach number M.

DUCT EFFICIENCY • The duct pressure efficiency ratio is defined as the ability of the duct to convert the kinetic or dynamic pressure energy at the inlet of the duct to the static pressure energy at the inlet of the compressor without a loss in total pressure . • It is in order of 98% if there is less friction loss.

RAM RECOVERY POINT • The Ram Recovery Point is that aircraft speed at which the ram pressure rise is equal to the friction pressure losses OR • That aircraft speed at which the compressor inlet total pressure is equal to the outside ambient air pressure. • A good subsonic duct has aircraft speed of 257.4 km/h for a good ram recovery point.

Supersonic Inlets • Even for supersonic flight it remains necessary that the flow leaving the inlet system be subsonic. It is required to have some means to • decelerate supersonic flow to subsonic speeds tolerable by existing compressors or fans.

Types of Supersonic Inlets •

•

•

Reverse Nozzle Diff or Converging Diverging Intakes Normal Shock Diff or Pitot Inlet Oblique Shock Diff

Reverse Nozzle Diff or Diffs with internal contraction or Converging Diverging Intakes • Deceleration from supersonic to subsonic flow speeds can be done by a simple normal shock with small stagnation pressure loss if the upstream Mach number is close to 1. • For high Mach number the loss across a single normal shock would be excessive. • Therefore it is better to use a combination of oblique shocks.

Normal-Shock diff • All existing compressors and fans require subsonic flow at their inlet with 0.5 < M2 < 0.8 at high power conditions. • So the inlet must reduce the flow Mach number from Mo > 1 to M2 < 1. • The simplest way to do this is with a Normal Shock. Prandtl – Meyer Relation for the normal • shock in a perfect gas is • V 1V 2 = a*2 = 2a0 2 / ᵞ + 1

•

M1 * M2 * = 1

Normal-Shock diff

Normal-Shock diff • For low supersonic speeds, such diffs are adequate because the stagnation pressure loss is small, but at Mo = 2, pt2 / pto ≈ 0.71, a serious penalty, and at Mo = 3 pt2 / pto ≈ 0.32. • For example the F-16 fighter has a simple normal shock diff, while the F-15 has an oblique shock diff.

Oblique - Shock diffs • The losses can be greatly reduced by decelerating the flow through one or more oblique shocks, the deflection and the pressure rise of each being small enough to be in the range where the stagnation pressure ratio is close to unity. • It is very important to understand that an Oblique Shock is in fact just a normal shock standing at an angle to the flow.

Oblique - Shock diffs

Oblique - Shock diffs • M1n is given in of Mon by the same relation given for M1 as a function of Mo. But Mon can be made close to 1. The condition for a weak sound wave is just • Mon = 1,

Oblique - Shock diffs • By choosing the wedge angle (or deflection angle) ∂ we can set the shock angle. • A series of weak oblique shocks, for each of which the Mn is near unity, hence all lying in the range of small pt loss, can yield an efficient diff.

SUPERSONIC INLETS WITH VARIABLE GEOMETRY

• This would work at one design Mach number, the one for which the isentropic area ratio between the incoming supersonic flow and the sonic throat is exactly the as-built area ratio A1 / Athroat. • But during the acceleration to this Mach number the fully supersonic flow cannot be established in the inlet without varying the geometry. • Imagine the inlet flying at M0 lower than the design Mach number. • The flow will look as depicted in the top right in

diagram shown in next slide.

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

STARTING THE DIFF • If the flow arrives undisturbed at the inlet, it could only occupy a fraction of it, the rest of the flow into the frontal area A1 is required to be disposed of which is called “Spillage”. This “Spillage” is accomplished by the • detached normal shock; behind it the flow is subsonic and it can turn around the inlet.

• The shock at the full flight Mach number is very lossy, and it is not practical to simply force the plane to continue accelerating to the design condition (there may not even be enough thrust left to do it). • What can be done is to manipulate the geometry to swallow the shock and reduce its strength. This is called “STARTING” THE DIFF.

STARTING OF THE DIFF

• To "START" THE DIFF, means to the shock through the convergent portion, there should be an increase in the throat area until the normal shock is just at the lip. • At that point, any further small increase in throat area causes the shock to jump rapidly to a position in the divergent part of the nozzle where the area is again A1. • This rapid jumping of shock from converging portion to diverging portion takes place because the shock is unstable in the converging section, but stable in the divergent section.

• This is accomplished by the flow due to which repositioning of the shock to a location nearer the throat, on the supersonic side takes place. • The process can continue until the shock is almost at the throat. • This repositioning of shock in throat on the supersonic side is called “STARTING OF THE DIFF”. For this successive step of acceleration 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 the upstream capture area Aa is determined by conditions downstream of the inlet.

Condition (b) • Low subsonic speed operation. Inlet is choked. • • The inlet mass flow rate is limited by the choking condition at At . * Since the flow is isentropic, A = A and the t • upstream capture area Aa + is given by

Condition (c) to (f) • In condition (c) to (f), the inlet flow velocity is increased gradually to design Mach No. MD to position the shock first in front of the inlet, then in the cowl or inlet lip, then it enters in the converging part and jumps rapidly to diverging part. • When the air intake starts operating in design Mach No. MD, the shock repositions itself to the throat 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

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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.

Modes of Inlet Operation SUPER - CRITICAL INLET OPERATION

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

Modes of Inlet Operation

FLOW INSTABILITY BUZZ • 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.

Variable Geometry Duct At 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 (2dimensional 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 a bellmouth inlet. It is a subsonic inlet. This type of inlet is essentially a bell shaped • funnel having carefully rounded shoulders which offers practically no air resistance. • The duct loss is so small that it is considered zero and all en e performance data can be gathered without any correction for inlet duct loss being necessary.

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