1) Jet engine inlets must supply the engine with airflow at high pressure to maximize thrust output. Inlet design is critical for both subsonic and supersonic aircraft. 2) For subsonic aircraft, inlets use either internal or external compression via divergent ducts to decelerate airflow without strong shockwaves. Supersonic inlets use convergent-divergent ducts with oblique shocks to decelerate airflow. 3) Proper inlet design considers boundary layer growth, external vs internal deceleration tradeoffs, and maintains high duct efficiency across a range of speeds and conditions. Inlet performance is measured by parameters like isentropic efficiency and stagnation pressure ratio.

1. SUBSONIC AND
SUPERSONIC AIR INTAKES
(JET PROPULSION)
SANJAY SINGH
Asst. Prof. and Head
Department Of Aeronautical Engineering
VMKV Engineering College
Salem (Tamilnadu)
sansiaf@gmail.com

2. 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.

3. • 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.

4. • 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.

5. • 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.

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

7. Internal Compression
Subsonic Intakes
• A divergent duct acts as a subsonic internal
compression diffuser.
• The pressure gradients of this 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 diffuser and hence
the associated friction loss.

8. Internal Compression
Subsonic Intakes

9. 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.

10. 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.

11. External Compression
Subsonic Intakes

12. SUBSONIC DUCTS

13. BOUNDARY LAYER

14. 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.

15. 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.

16. 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.

17. 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.

18. Thrust on inlet Surface

19. Thrust on inlet Surface
• Net momentum flux out of the control
volume is
• Net momentum flux is
• Bernoulli’s law

20. The above relation shows that the
greater external deceleration (i.e.
the smaller the ratio ui / ua ), the
larger must be thrust increment.

21. Coefficient of Pressure (Cp)
• 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.

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

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

24. 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
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.
ui / ua > 0.8), its effect on minimum nacelle
size is quite small.

25. 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.

26. 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
diffuser 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.

28. Inlet Performance Criterion

29. Performance Criterion
1. Isentropic Efficiency of a Diffuser (defined
in terms of temperature rise).
State 02s is defined as the state that would be
reached by isentropic compression to the actual
outlet stagnation pressure.

30. • Since,
• Diffuser efficiency
can be written as

31. Ram Efficiency
2. Ram Efficiency (Defined in terms of
pressure rises)
ᶯr = (P02 - Pa ) / P0a - Pa

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

33. • Diffuser 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.

34. 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.

35. 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.

37. 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.

38. Types of Supersonic Inlets
• Reverse Nozzle Diffuser or
Converging
Diverging
Intakes
• Normal Shock Diffuser or Pitot
Inlet
• Oblique Shock Diffuser

39. Reverse Nozzle Diffuser or Diffusers
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.

41. Normal-Shock diffuser
• 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

42. Normal-Shock diffuser

44. Normal-Shock diffuser
• For low supersonic speeds, such diffusers
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 diffuser, while the F-15 has
an oblique shock diffuser.

45. Oblique - Shock diffusers
• 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.

46. Oblique - Shock diffusers

47. Oblique - Shock diffusers
• M1n is given in terms 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,

48. Oblique - Shock diffusers
• 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 diffuser.

49. SUPERSONIC INLETS WITH
VARIABLE GEOMETRY

50. • 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.

51. 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.

52. STARTING THE DIFFUSER
• 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.

53. • 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
DIFFUSER.

54. STARTING OF THE DIFFUSER

55. • To "START" THE DIFFUSER, means to pass
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.

56. • 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
DIFFUSER” . For this successive steps of
acceleration is followed.

57. Successive Steps in
Acceleration of a CD inlet

58. 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.

59. 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, At = A* and the
upstream capture area Aa + is given by

60. 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.

61. 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.

62. 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.

63. Modes of Inlet 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.

64. Modes of Inlet Operation

65. 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.

66. 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.

67. 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 engine performance data can be
gathered without any correction for inlet duct
loss being necessary.

68. THANK YOU.