Components of jet engines

Radiusing of the lip prevents flow separation and compressor inlet distortion at low speeds during crosswind operation and take-off rotation.

There are basically two forms of shock waves: A sharp-lipped version of the pitot intake, described above for subsonic applications, performs quite well at moderate supersonic flight speeds.

However, below the shock-on-lip flight Mach number, the shock wave angle/s are less oblique, causing the streamline approaching the lip to be deflected by the presence of the cone/ramp.

This discontinuity is overcome by the normal shock moving to a lower cross-sectional area in the ducting, to decrease the Mach number at entry to the shockwave.

The intake on the SR-71 had a translating conical spike which controlled the shock wave positions to give maximum pressure recovery.

Two vertical ramps were used in the F-4 Phantom intake, the first with a fixed wedge angle of 10 degrees and the second with a variable additional deflection above Mach 1.2.

Many compressors are fitted with anti-stall systems in the form of bleed bands or variable geometry stators to decrease the likelihood of surge.

This implies a higher high-pressure shaft speed, to maintain the datum blade tip Mach number on the rear compressor stage.

Stress considerations, however, may limit the shaft speed increase, causing the original compressor to throttle-back aerodynamically to a lower pressure ratio than datum.

[16] Since the turbine cannot withstand the stoichiometric temperatures (a mixture ratio of around 15:1) in the combustion zone, the compressor air remaining after supplying the primary zone and wall-cooling film, and known as dilution air, is used to reduce the gas temperature at entry to the turbine to an acceptable level (an overall mixture ratio of between 45:1 and 130:1 is used[17]).

The turbine needs fewer stages than the compressor, mainly because the higher inlet temperature reduces the deltaT/T (and thereby the pressure ratio) of the expansion process.

All jet engines require high temperature gas for good efficiency, typically achieved by combusting hydrocarbon or hydrogen fuel.

After removing heat from the blade material, the air (now fairly hot) is vented, via cooling holes, into the main gas stream.

Another solution is to use an ultra-efficient turbine rim seal to pressurize the area where the cooling air passes across to the rotating disc.

This is done using primary and secondary airholes which allow a thin layer of air to cover the inner walls of the chamber preventing excessive heating.

A bypass tube is a component, usually found on turbojets that is an external pipe/s that channels air from the compressor to the propelling nozzle and afterburner section.

[19] Bypass tubes only work best at supersonic speeds from mach 2+ when the ram air effect is more than what the compressor can handle where the turbojet essentially becomes a ramjet.

[citation needed] Notable examples using bypass tubes are the Brandner E-300, Rolls-Royce Olympus and the Pratt & Whitney J58 the most common.

The fuel must be reduced with altitude to maintain the same a/f ratio as the lower ambient pressure means less weight of air entering the engine (early engine controls used a Barostat or Barometric Pressure Control depending on the type of fuel pump, fixed or variable displacement).

At high engine speeds over-speeding and over-temperaturing (going beyond the maximum allowable) must be prevented to avoid turbine blade damage.

Others, such as the HP rotor speed, will modify the pilot's request as necessary before sending a signal to the torque motor which sets the position of the FMV.

The HMU also sends fuel hydraulic signals using FADEC-controlled individual torque motors to actuators for the variable stator vanes, low and high pressure turbine clearance control, high pressure compressor clearance control and a motor for the variable bleed valves.

This is an impulse turbine impacted by burning gases from a cartridge, usually created by igniting a solid propellant similar to gunpowder.

However, the turbine is turned by burning gases - usually the fuel is isopropyl nitrate (or sometimes Hydrazine) stored in a tank and sprayed into a combustion chamber.

Most commercial aircraft and large military transport airplanes usually use what is called an auxiliary power unit (APU).

This is the typical location for an APU on most commercial airliners although some may be within the wing root (Boeing 727) or the aft fuselage (DC-9/MD80) as examples and some military transports carry their APUs in one of the main landing gear pods (C-141).

A variation of this is the APU installed in a Boeing F/A-18 Hornet; it is started by a hydraulic motor, which itself receives energy stored in an accumulator.

A nose bullet is a component of a turbojet used to divert air into the intake, in front of the accessory drive and to house the starter motor.

Depending on various conditions, such as flying through heavy rainfall, the igniter continues to provide sparks to prevent combustion from failing if the flame inside goes out.

As a result, some PRVs can adjust their spring force values using this pressure change in the bearing chamber proportionally to keep the lubricant flow constant.

Diagram of a typical gas turbine jet engine. Air is compressed by the compressor blades as it enters the engine, and it is mixed and burned with fuel in the combustion section. The hot exhaust gases provide forward thrust and turn the turbines which drive the compressor blades. 1. Intake
2. Low pressure compression
3. High pressure compression
4. Combustion
5. Exhaust
6. Hot section
7. Turbines Low and High pressure
8. Combustion chambers
9. Cold section
10. Air inlet
Basic components of a jet engine (Axial flow design)
A blade with internal cooling as applied in the high-pressure turbine
Pitot intake operating modes
Thin rounded intake lip
An unswept lip generate a shock wave, which is reflected multiple times in the inlet. The more reflections before the flow gets subsonic, the better pressure recovery
Axial compressor animation showing rotating blades and stationary stators
The 17-stage axial compressor on the General Electric J79 . Air moves from left to right. Each compressor stage is a row of rotor blades which give the air tangential velocity followed by a stationary row of stator blades which slow the air and raise its static pressure. The top casing holding the stators has been removed to show the rotor blades. Visible on the bottom casing are the actuating levers for the six stages of variable stators. The blades and air passage get smaller from left to right because the volume of each pound of air gets smaller as the air is compressed from 1 atmosphere to 12.
Combustion liners in GE J79. The top outer combustion casing has been removed. Air from the compressor moves from left to right. Some enters the liner at the left where the fuel nozzle is located. The remaining air passes along the outside of the liner and enters the holes downstream of the primary zone to complete the combustion and dilute combustion products to a temperature acceptable to the turbine first stage nozzle guide vanes visible at the right. Two of the liners have spark plugs and during a start the remaining eight are ignited through cross-fire tubes which link all the liners. One tube is just visible between the lower and middle liners.
The 3-stage turbine on the GE J79. Each stage is a stationary ring of nozzle guide vanes followed by spinning blades. The gas is moving from left to right and the 2nd and 3rd vane rings have been removed to better show the blades. The first ring shows the shape of the vanes and how they turn the gas from the combustor into a tangential direction necessary to spin the bladed disc.