Turbojet

For the transportation company in southern China, see TurboJET.

For fuller treatment of this class of engine, see Jet engine

Turbojets are the simplest and oldest kind of general purpose jet engine. Two different engineers, Frank Whittle in Britain and Hans von Ohain in Germany, developed the concept during the late 1930s. Fighter aircraft, fitted with turbojet engines, first entered service in 1944, towards the end of World War II.

A turbojet engine is used primarily to propel aircraft. Air is drawn into the rotating compressor via the intake and is compressed to a higher pressure before entering the combustion chamber. Fuel is mixed with the compressed air and ignited by flame in the eddy of a flame holder. This combustion process significantly raises the temperature of the gas. Hot combustion products leaving the combustor expand through the turbine, where power is extracted to drive the compressor. Although this expansion process reduces the turbine exit gas temperature and pressure, both parameters are usually still well above ambient conditions. The gas stream exiting the turbine expands to ambient pressure via the propelling nozzle, producing a high velocity jet in the exhaust plume. If the jet velocity exceeds the aircraft flight velocity, there is a net forward thrust upon the airframe.

Modern jet engines are mainly turbofans, where some (if not most) of the air entering the intake bypasses the combustor.

Although ramjet engines are simpler in design (virtually no moving parts) they are incapable of operating at low flight speeds.

Air intake

Preceding the compressor is the air intake (or inlet), which is designed to recover, as efficiently as possible, the ram pressure of the streamtube approaching the intake. Downstream of the intake, air enters the compression system.

Compressor

The compressor, which rotates at very high speed, adds energy to the airflow, at the same time squeezing it into a smaller space, thereby increasing its pressure and temperature.

In most turbojet-powered aircraft, bleed air is extracted from the compressor section at various stages to perform a variety of jobs including air conditioning/pressurization, engine inlet anti-icing, and many others.

Several types of compressor are used in turbojets and gas turbines in general: axial, centrifugal, axial-centrifugal, double-centrifugal, etc.

Early turbojet compressors had overall pressure ratios as low as 5:1 (as do a lot of simple auxiliary power units and small propulsion turbojets today). Aerodynamic improvements, plus splitting the compression system into two separate units and/or fitting anti-stall systems, enabled later turbojets to have overall pressure ratios of 15:1 or more. In comparison, modern civil turbofan engines have overall pressure ratios as high as 44:1 or more.

After leaving the compressor section, the compressed air enters the combustor.

Combustor

The burning process in the combustor is significantly different from that in a piston engine. In a piston engine the burning gases are confined to a small volume and, as the fuel burns, the pressure increases dramatically. In a turbojet the air and fuel mixture passes, unconfined, through the combustion chamber. As the mixture burns its temperature increases dramatically, the pressure actually decreasing a few percent. In detail, the fuel-air mixture must be brought almost to a stop so that a stable flame can be maintained, this occurs just after the beginning of the combustion chamber. The aft part of this flame front is allowed to progress rearward in the engine. This ensures that the rest of the fuel is burned as the flame becomes hotter when it leans out, and because of the shape of the combustion chamber the flow is accelerated rearwards. Some pressure drop is unavoidable, as it is the reason why the expanding gases travel out the rear of the engine rather than out the front. Less than 25% of the air is involved in combustion, in some engines as little as 12%, the rest acting as a reservoir to soak up the heating effect of the fuel burning.

Another difference between piston engines and jet engines is that the peak flame temperature in a piston engine is experienced only momentarily, and for a small portion of the entire cycle. The combustor in a jet engine is exposed to the peak flame temperature continuously and operates at a pressure high enough that a stoichiometric fuel-air ratio would melt the can and everything downstream. Instead, jet engines run a very lean mixture, so lean that it would not normally support combustion. A central core of the flow (primary airflow) is mixed with enough fuel to burn readily. The cans are carefully shaped to maintain a layer of fresh unburned air between the metal surfaces and the central core. This unburned air (secondary airflow) mixes into the burned gases to bring the temperature down to something the turbine can tolerate.

Turbine

Hot gases leaving the combustor are allowed to expand through the turbine. In the first stage the turbine is largely a reaction turbine (similar to a pelton wheel) and rotates because of the impact of the hot gas stream. Later stages are convergent ducts that accelerate the gas rearward and gain energy from that process. Pressure drops, and energy is transferred into the shaft. The turbine's rotational energy is used primarily to drive the compressor. Some shaft power is extracted to drive accessories, like fuel, oil, and hydraulic pumps. Because of its significantly higher entry temperature, the turbine pressure ratio is much lower than that of the compressor. In a turbojet almost two thirds of all the power generated by burning fuel is used by the compressor to compress the air for the engine.

Nozzle

After the turbine, the gases are allowed to expand through the exhaust nozzle to atmospheric pressure, producing a high velocity jet in the exhaust plume. In a convergent nozzle, the ducting narrows progressively to a throat. The nozzle pressure ratio on a turbojet is usually high enough for the expanding gases to reach Mach 1.0 and choke the throat. Normally the flow will go supersonic in the exhaust plume, external to the engine.

If, however, a convergent-divergent "de Laval" nozzle is fitted, the divergent (increasing flow area) section allows the gases to reach supersonic velocity within the nozzle itself. This is slightly more efficient on thrust, than using a convergent nozzle. There is, however, the added weight and complexity, since the con-di nozzle must be fully variable, to cope basically with engine throttling.

Net thrust

Below is an approximate equation for calculating the net thrust of a turbojet:

$F_n = m \cdot (V_{jfe} - V_a)$

where:

$m = \,$ intake mass flow

$V_{jfe} =\,$ fully expanded jet velocity (in the exhaust plume)

$V_a =\,$ aircraft flight velocity

Whilst the $m.V_{jfe}\,$ term represents the nozzle gross thrust, the $m.V_a\,$ term represents the ram drag of the intake. Obviously, the jet velocity must exceed that of the flight velocity if there is to be a net forward thrust on the airframe.

Thrust to power ratio

A simple turbojet engine will produce thrust of approximately: 2.5 pounds force per horsepower (15 mN/W).

Cycle improvements

Increasing the overall pressure ratio of the compression system raises the combustor entry temperature. Therefore, at a fixed fuel flow and airflow, there is an increase in turbine inlet temperature. Although the higher temperature rise across the compression system, implies a larger temperature drop over the turbine system, the nozzle temperature is unaffected, because the same amount of heat is being added to the system. There is, however, a rise in nozzle pressure, because overall pressure ratio increases faster than the turbine expansion ratio. Consequently, net thrust increases, whilst specific fuel consumption (fuel flow/net thrust) decreases.

So turbojets can be made more fuel efficient by raising overall pressure ratio and turbine inlet temperature in unison. However, better turbine materials and/or improved vane/blade cooling are required to cope with increases in both turbine inlet temperature and compressor delivery temperature. Increasing the latter requires better compressor materials..

Early designs

Early German engines had serious problems controlling the turbine inlet temperature. Their early engines averaged only ten hours of operation before failing—often with chunks of metal flying out the back of the engine when the turbine overheated. British engines tended to fare better due to better metals. The Americans had the best materials because of their reliance on turbosupercharging in high altitude bombers of World War II. For a time some US jet engines included the ability to inject water into the engine to cool the compressed flow before combustion, usually during takeoff. The water would tend to prevent complete combustion and as a result the engine ran cooler again, but the planes would takeoff leaving a huge plume of smoke.

Today these problems are much better handled, but temperature still limits airspeeds in supersonic flight. At the very highest speeds, the compression of the intake air raises the temperature to the point that the compressor blades will melt. At lower speeds, better materials have increased the critical temperature, and automatic fuel management controls have made it nearly impossible to overheat the engine.

Sources

Constructing A Turbocharger Turbojet Engine. Edwin H. Springer. Turbojet Technologies 2001.

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