In this article we will discuss about:- 1. Turbojet Engine 2. Turbofan Engine 3. Turboprop Engine.
The turbojet engine consists of a diffuser at the entrance which slows down the entrance air and thereby compresses it, called the ramming effect; a simple open gas turbine cycle and an exit nozzle which expands the gas and converts the thermal energy of the exit gas into kinetic energy which is emitted out in the form of a jet. The increased velocity of air thereby produces thrust.
A schematic arrangement of a turbo-jet engine is shown in Fig. 35.1. A corresponding T-S diagram is shown. The basic cycle for the turbojet engine is the joule or Brayton cycle. In process 0 to 1, the air entering from the atmosphere is diffused isentropically from velocity Va down to zero i.e., V1 = 0. This means that the diffuser has an efficiency of 100 percent. This is termed as ram-compression. Process 0-1′ is the actual compression (ram).
Process 1′-2 shows the isentropic compression of air and 1’—2 shows the actual compression of air. 2′-3 shows the heat addition at constant pressure p2 = p3. Process 3-4 shows isentropic expansion of gas in the turbine and 3-4′ shows the actual expansion in the turbine. Process 4-5 shows the isentropic expansion of the gas in the nozzle and 4’—5′ shows the actual expansion of gas in the nozzle. Consider 1 kg of working fluid flowing through the system.
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 independently during the late 1930s. On 27 August 1939 the Heinkel He 178 became the world’s first aircraft to fly under turbojet power, thus becoming the first practical jet plane.
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 momentum of the exhaust stream exceeds the momentum of the intake stream, the impulse is positive, thus, there is a net forward thrust upon the airframe.
Figure 35.5 shows the P-V diagram of turbojet engine and Fig. 35.6 is the T-S diagram of the turbojet engine.
While analysing the cycle we assume that:
1. Specific heat is constant.
2. There is no pressure drop in combustion chamber.
0 – 1 is the actual compression (ramming) in the diffuser with a diffuser efficiency of ηd.
1 – 2 is the actual compression in the compressor with the compression efficiency ηc.
2- 3 Heating or heat addition in the combustion chamber.
3- 4 Actual expansion of gases in gas-turbine with an isentropic efficiency of ηt.
4- 5 Actual expansion of gases through jet-nozzle with an nozzle efficiency of ηn.
The thrust developed by the engine should overcome the drag on the air-craft and in doing so develops power, which is called thrust power and is given by-
Specific thrust is a criterion of the size of the engine required to produce a given total thrust. From the equation for the thrust (2a) we observe that if the effective jet velocity Ve is kept constant, then the specific thrust decreases with increase in plane velocity Va = Vp or flight speed. Specific thrust will be maximum when a = 0 i.e., Va = 0 or flight velocity is zero.
A turbofan is a type of airplane engine, similar to a turbojet. It essentially consists of a ducted fan with a smaller diameter turbojet engine mounted behind it that powers the fan. Part of the airstream from the ducted fan passes through the turbojet where it is burnt to power the fan, but the majority of the flow bypasses it, and produces most of the thrust.
All of the jet-engines that power currently-manufactured commercial jet aircraft are turbofans. They are mainly used commercially because they are highly efficient and relatively quiet in operation.
After World War II, 2-spool (shaft) turbojets were developed to make it easier to throttle-back compression systems with a high design overall pressure ratio (i.e., combustor inlet pressure/intake delivery pressure).
Adopting the 2-spool arrangement enables the compression system to be split into two, with a Low Pressure (LP) Compressor supercharging a High Pressure (HP) Compressor. Each compressor is mounted on a separate co-axial shaft, driven by its own turbine. Otherwise a 2-spool turbojet is much like a single spool engine.
Modern turbofans evolved from the 2-spool axial-flow turbojet engine, essentially by increasing the relative size of the Low Pressure (LP) Compressor to the point where some of the air exiting the unit actually bypasses the core (or gas generator) stream, passing through the main combustor. This bypass air either expands through a separate propelling nozzle, or is mixed with the hot gases leaving the Low Pressure (LP) Turbine, before expanding through a Mixed Stream Propelling Nozzle.
Owing to a lower jet velocity, a modern civil turbofan is quieter than the equivalent turbojet. Turbofans also have a better thermal efficiency. In a turbofan, the LP Compressor is often called a fan. Civil turbofans usually have a single fan stage, whereas most military turbofans have multi-stage fans.
It is a parameter often used for classifying turbofans, defined as ratio of bypassed airflow to combustor airflow.
The noise of any type of jet engine is strongly related to the velocity of the exhaust gases. High bypass ratio (i.e., low specific thrust) turbofans are relatively quiet compared to turbojets and low bypass ratio (i.e., high specific thrust) turbofans.
A low specific thrust engine has a low jet velocity by definition, as the following approximate equation for net thrust implies:
So for zero flight velocity, specific thrust is directly proportional to jet velocity. Relatively speaking, low specific thrust engines are large in diameter to accommodate the high airflow required for a given thrust.
A high specific thrust/low bypass ratio turbofan normally has a multi-stage fan, developing a relatively high pressure ratio and, thus, yielding a high (mixed or cold) exhaust velocity. The core airflow needs to be large enough to give sufficient core power to drive the fan. A smaller core flow/higher bypass ratio cycle can be achieved by raising the (HP) turbine rotor inlet temperature.
A bypass flow can only be introduced if the turbine inlet temperature is allowed to increase, to compensate for a correspondingly smaller core flow. Improvements in turbine cooling/material technology would facilitate the use of a higher turbine inlet temperature, despite increases in cooling air temperature, resulting from a probable increase in overall pressure ratio.
Efficiently done, the resulting turbofan would probably operate at a higher nozzle pressure ratio than the turbojet, but with a lower exhaust temperature to retain net thrust. Since the temperature rise across the whole engine (intake to nozzle) would be lower, the (dry power) fuel flow would also be reduced, resulting in a better specific fuel consumption (SFC).
Since the 1970s, most jet fighter engines have been low/medium bypass turbofans with a mixed exhaust, afterburner and variable area final nozzle—the first afterburning turbofan was the Pratt and Whitney TF30.
Unlike the main combustor, where the integrity of the downstream turbine blades must be preserved, an afterburner can operate at the ideal maximum temperature (-i.e., about 2100 K). Now, at a fixed total applied fuel: air ratio, the total fuel flow for a given fan airflow will be the same, regardless of the dry specific thrust of the engine.
However, a high specific thrust turbofan will, by definition, have a higher nozzle pressure ratio, resulting in a higher afterburning net thrust and, therefore, a lower afterburning specific fuel consumption. However, high specific thrust engines have a high dry SFC. The situation is reversed for a medium specific thrust afterburning turbofan: i.e., poor afterburning SFC/good dry SFC.
The former engine is suitable for a combat aircraft which must remain in afterburning combat for a fairly long period, but only has to fight fairly close to the airfield (i.e., cross border skirmishes). The latter engine is better for an aircraft that has to fly some distance, or loiter for a long time, before going into combat. However, the pilot can only afford to stay in afterburning for a short period, before his/her fuel reserves become dangerously low.
Modern low-bypass military turbofans include the Pratt and Whitney F119, the Eurojet EJ200 and the General Electric F110, all of which feature a mixed exhaust, afterburner and variable area propelling nozzle. Non-afterburning engines include the Rolls-Royce/Turbomeca Adour (afterburning in the SEPECAT Jaguar) and the unmixed, vectored thrust. Rolls-Royce Pegasus.
The low specific thrust/high bypass ratio turbofans used in today’s civil jetliners (and some military transport aircraft) evolved from the high specific thrust/low bypass ratio turbofans used in such aircraft back in the 60’s.
Low specific thrust is achieved by replacing the multi-stage fan with a single stage unit. Unlike some military engines, modern civil turbofans do not have any stationary inlet guide vanes in front of the fan rotor. The fan is scaled to achieve the desired net thrust.
The core (or gas generator) of the engine must generate sufficient Core Power to at least drive the fan at its design flow and pressure ratio. Through improvements in turbine cooling/material technology, a higher (HP) turbine rotor inlet temperature can be used, thus facilitating a smaller (and lighter) core and (potentially) improving the core thermal efficiency.
Reducing the core mass flow tends to increase the load on the LP turbine, so this unit may require additional stages to reduce the average stage loading and to maintain LP turbine efficiency. Reducing core flow also increases bypass ratio (5: 1, or more, is now common).
Further improvements in core thermal efficiency can be achieved by raising the overall pressure ratio of the core. Improved blade aerodynamics reduces the number of extra compressor stages required. With multiple compressors (i.e., LPC, IPC, HPC) dramatic increases in overall pressure ratio have become possible. Variable geometry (i.e., stators) enable high pressure ratio compressors to work surge-free at all throttle settings.
The first high-bypass turbofan engine was the General Electric TF39, built to power the Lockheed C-5 Galaxy military transport aircraft. The civil General Electric CF6 engine used a derived design. Other high-bypass turbofans are the Pratt and Whitney JT9D, the three-shaft Rolls-Royce RB211 and the CFM International CFM56. More recent large high-bypass turbofans include the Pratt and Whitney PW4000, the three-shaft Rolls-Royce Trent, the General Electric GE90, and the General Electric GEnx.
The significantly higher thrust provided by high-bypass turbofan engines also made civil wide-body aircraft practical and economical. In addition to the vastly increased thrust, these engines are also generally quieter.
This is not so much due to the higher bypass ratio, as to the use of low pressure ratio, single stage, fans, which significantly reduce specific thrust and, thereby, jet velocity. The combination of a higher overall pressure ratio and turbine inlet temperature improves thermal efficiency. This, together with a lower specific thrust (better propulsive efficiency), leads to a lower specific fuel consumption.
For reasons of fuel economy, and also of reduced noise, almost all of today’s jet airliners are powered by high- bypass turbofans. Although modern military aircraft tend to use low bypass ratio turbofans, military transport aircraft (e.g. C-17) mainly use high bypass ratio turbofans (or turboprops) for fuel efficiency.
Turbofan engines come in a variety of engine configurations.
This is probably the simplest configuration, comprising a fan and high pressure compressor driven by a single turbine unit, all on the same shaft. Despite the simplicity of the turbomachinery configuration, it requires a variable area mixer to facilitate part-throttle operation.
One of the earliest turbofans was a derivative of the General Electric J79 turbojet, known as the CJ805, which featured an integrated aft fan/low pressure (LP) turbine unit located in the turbojet exhaust jetpipe. Hot gas from the turbojet turbine exhaust expanded through the LP turbine, the fan blades being a radial extension of the turbine blades. One of the problems with the Aft Fan configuration is hot gas leakage from the LP turbine to the fan.
Many turbofans have the Basic Two Spool configuration where both the fan and LP turbine are mounted on a second (LP) shaft, running concentrically with the HP spool. At the smaller thrust sizes, instead of all-axial blading, the HP compressor configuration may be axial-centrifugal, double-centrifugal or even diagonal/centrifugal.
Higher overall pressure ratios are achieved by either raising the HP compressor pressure ratio or adding an Intermediate Pressure (IP) Compressor between the fan and HP compressor, to supercharge or boost the latter unit.
The IP compressor is mounted on a separate (IP) shaft, running concentrically with the LP and HP shafts, and is driven by a separate Intermediate Pressure (IP) Turbine.
As bypass ratio increases, the mean radius ratio of the fan and LP turbine increases. Consequently, if the fan is to rotate at its optimum blade speed the LP turbine blading will run slow, so additional LPT stages will be required, to extract sufficient energy to drive the fan. Introducing a reduction gearbox, with a suitable gear ratio, between the LP shaft and the fan, enables both the fan and LP turbine to operate at their optimum speeds. This is not a popular solution, since high power gearboxes tend to be unreliable.
The advantage of turbofan or bypass engine is that, at lower jet velocities, it gives better propulsive efficiency and fuel economy compared with a turbojet engine for a given thrust, as less kinetic energy is wasted to atmosphere.
This applies particularly at aircraft speeds below sonic velocity and for long range aircraft. The bypass air helps for fire protection by the cool air in the bypass which surrounds the hot parts of the engine. Such engines are mostly used on jumbo-jets for civil aviation.
A Turboprop engine is a type of gas turbine engine which uses most of its power to drive a propeller. The propeller of a turboprop is very similar to that used by piston or reciprocating engines, but turboprops usually use a constant velocity propeller. Turboprop engines are generally used on small or slow subsonic aircraft, but some aircraft outfitted with turboprops have cruising speeds.
A turboprop consists of an intake, compressor, combustor, turbine and a propelling nozzle. Air is drawn into the intake and compressed by the compressor. Fuel is then added to the compressed air in the combustor. The hot combustion gases expand through the turbine, to provide power to the turbine.
Part of this power goes to the compressor, and the rest, through the reduction gearing, to the propeller. Further expansion of the gases occurs in the propelling nozzle, where the gasses are expelled, providing approximately 25% of the thrust.
A turboprop engine is similar to a turbojet, but has additional fan blades in the turbine stage to recover more power from the engine to turn the propeller.
In a turboprop much of the jet thrust is sacrificed in favour of shaft power, which is obtained by extracting additional power (to that necessary to drive the compressor) from the turbine expansion process. While the power turbine may be integral with the gas generator section, many turboprops today feature a Free Power Turbine, on a separate coaxial shaft.
This enables the propeller to rotate freely, independent of compressor speed. Owing to the additional expansion in the turbine system, the residual energy in the jet is fairly low (<10% of total thrust, including that of the propeller).
Because the propeller is very much larger in diameter than the power turbine, the tip speed of the propeller can become supersonic. Consequently, to prevent this, a speed reduction gearbox is inserted between the power turbine and propeller shafts. The gearbox is part of the engine, whereas in a turbo-shaft the (helicopter) rotor reduction gearbox is remote from the engine.
Residual thrust on a turbo-shaft is avoided by further expansion in the turbine system and/or truncating and turning the exhaust through 90 degrees, to produce two opposing jets.
Apart from the above, there is very little difference between a turboprop and a turbo-shaft.
Turboprops are very efficient at modest flight speeds (below 724 km/h or 450 mph), because the jet velocity of the propeller (and exhaust) is relatively low. Consequently, small commuter aircraft and military transports tend to feature turboprop engines.
Although turboprops are used in some General Aviation applications, their high price deters more widespread acceptance except for very high performance STOL applications. One of the most common turboprop military aircraft is the C-130 Hercules transport.
While most modern turbojet and turbofan engines use axial-flow compressors, turboprop engines usually contain at least one stage of centrifugal compression, because of the small size of the engines.
Propellers lose efficiency as aircraft speed increases, which is why turboprops are not used on higher-speed aircraft.