The idea of utilizing the physical principle of reaction on a large scale by means of rockets is usually attributed to China in the thirteenth century. Not until after the Second World War, however, did rocket technology mature to a state which made the idea of space travel a practical possibility, owing largely to a giant step forward during the war itself.
Although the earliest models of the steam turbine date back as early as the 17th century, practical applications of the turbine engine had to wait until the turn of the 20th century. Today, the gas turbine engine is the most widespread and most effective method of aircraft propulsion, having almost totally displaced the reciprocating engine, which, up to the 1960s, was the common power source in aviation.
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There are four types of engines I will be talking about; turbojet, turbofan, turboprop, or turboshaft. The gas turbine represents one of the most technological achievements in aviation, the successful introduction of which made possible a tremendous acceleration of progress in all fields of aviation.
Some of the historical milestones that are major steps toward turbine engine development, ending in the use of the gas turbine for aircraft propulsion are as follows:
1687 – the English philosopher and mathematician Sir Isaac Newton formulates three laws of motion which form the basis of modern jet propulsion, according to which:
1) a body remains either at rest or in the motion of constant velocity unless an external force acts on the body;
2) the sum of forces acting on a body equals the product of the body’s mass times acceleration produced by these forces ( i.e. force = mass times acceleration);
3) for every force acting on a body, the body exerts a force of equal magnitude and opposite direction along the same line as the original force.
1791 – John Barber, an Englishmen, was granted a patent for a gas-driven turbine engine which utilized the thermodynamic cycle of the modern gas turbine. The power plant was to be comprised of a gass generator with compressor, combustion chamber and a turbine wheel – components that are fundamental to today’s engine. This engine was never built.
1918 – Sanford Moss, at General Electric in the United States, develops an exhaust turbocharger for reciprocating areo-engines. This is the first application of a gas turbine in an aircraft propulsion system.
1930 – Frank Whittle applied for his patent in which he describes a jet engine with a multi-staged axial compressor followed by a centrifugal compressor, annular combustion chamber, single-stage axial turbine, and a nozzle. The patent was titled ‘Improvements relating to the Propulsion of Aircraft and other Vehicles’. The engine, which ran on liquid fuel, was successfully run in April of 1937.
These early steps laid the foundation of modern high-thrust engines.
Different types of engines exist, according to their tasks. Turbojet and turbofan engines provide propulsion forces directly by reaction forces generated by the exhaust gas. Turbofan engines, are classified according to the portion of mass airflow that is bypassed around the basic engine and are typically denoted as high-bypass or low-bypass-ratio engines.
The turbojet is made up of the following components:
- Multi-stage compressor
- Single or multi-stage turbine
In order to function properly to produce thrust, an air intake and an exhaust system are required to process the airflow. The air first enters the intake section, which must deliver a smooth stream of air for the compressor. The compressor’s job is to raise the pressure of the air. This also leads to a rise not only in pressure but also in temperature and density. On leaving the compressor, the pressurized air enters the combustion chamber, where fuel is injected and burned, thus adding more energy to the airflow. It is here that the airflow is processed to take on the characteristics of a gas. The first station where energy is absorbed from the gas is in the gas turbine.
The energy content of hot gas is not depleted when the gas discharges from the turbine. A major part of the heat and pressure energy of the gas is still available to be converted into kinetic energy. This is the task of the exhaust nozzle. High exhaust velocity is a prerequisite to the generation of thrust. Exhaust velocity may be increased further by afterburning, a fuel-exhausting method of adding more heat downstream of the turbine.
The turbofan has emerged as the most common type of gas turbine engine for aircraft propulsion. Like that of the turboprop, the turbine section is designed to absorb more energy than would be necessary to drive the compressor alone. The excess power is used to drive a fan, a low-pressure compressor of larger diameter arranged upstream of the main compressor. Part of the air entering the engine intake bypasses the inner engine and expands in a separate nozzle to provide cold rust.
The amount of air that is bypassed in relation to the air that passes through the core engine is turned bypassed-ratio. A distinction is made between low and high bypass ratio engines. Low bypass-ratio engines are used with supersonic combat aircraft. The high bypass-ratio engines are used with high subsonic military and commercial airliners.
In a turboprop, the energy of the hot gas is used to drive an additional, but separate turbine, which in turn provides shaft power to drive a propeller. The gas when exhausting from the nozzle has transmitted most of its energy to the turbines, which a small amount of energy remaining for the generation of thrust.
The basic layout of a turboprop is slightly different than a turbojet because it has an additional turbine to drive the propeller, a two-spool arrangement of the rotational machinery, and a mechanical reduction gear to convert the high rotational speed of the turbine to the more moderate speed of the propeller. A turboprop is designed to accelerate a high massed flow to a low velocity. This results in unsurpassed fuel efficiency, at the expense of flight speed and cabin noise.
In a turboshaft engine, all of the usable hot gas energy is extracted and converted into shaft power, by an additional (free) turbine. This type of engine is typically used with helicopters but is also employed in auxiliary power units to provide power for aircraft ground operation.
The compressor of a turboshaft engine first raises the pressure of the incoming air, which is then guided into the combustion chamber. After being mixed with vaporized fuel and burned, the hot gas expands completely through two separate turbines. The first turbine drives the compressor; the second delivers horsepower to drive the helicopter rotor. The gas is finally released through an exhaust duct without producing any thrust.
I will now tell you about the five major turbomachinery components of a jet engine: the air intake, compressor, combustion chamber, turbine and exhaust section. Each of these contributes uniquely to the generation of thrust.
In any application-subsonic transport or supersonic fighter-the, air intake is essentially a fluid flow duct whose task is to process the airflow in a way that ensures the engine functions properly to generate thrust.
The intake must be designed to provide the appropriate amount of airflow required by the engine and, furthermore, that this flow when leaving the intake section to enter the compressor will be uniform, stable, and of high quality. These conditions must be met not only during all phases of flight but on the ground, with the aircraft at rest and the engine demanding maximum thrust prior to take-off. Good intake design is therefore a prerequisite if installed engine performance is to come close to performance figures obtained at the static test bench.
Present-day turbine aero-engines require subsonic flow at the entry to the compressor, even if the aircraft is flying at supersonic speed. The task of the air intake is therefore to decelerate the supersonic external flow to a subsonic speed acceptable to the compressor.
In each category of a turbine-driven aero-engine-turbojet, turbofan, turboprop or turboshaft – the compressor is one of the most important components. It is the task of the compressor to increase the pressure of the airstreams that is furnished by the air intake. This process is accomplished by supplying mechanical energy (=work) to the compressor, the rotating blades of which exert aerodynamic forces on the airflow. At the compressor outlet, a stream of highly compressed air is discharged to the combustion chamber, where more energy is added in the form of heat.
In a compressor, mechanical energy is converted into pressure energy. The amount of energy required, and the quality of achieved energy conversion, is characterized by compressor performance parameters. The most important parameters are:
- compressor efficiency
- compressor total pressure ratio
- air-flow rate
The efficiency parameter denotes the amount of energy supplied to the compressor from the turbine, by means of the rotor shaft, which results in an increase of pressure energy. This parameter, therefore, denotes the amount of loss that is always incurred by converting energy.
Next, there is the compressor pressure ratio. This parameter is defined as the ratio of the total pressure at compressor discharge and at compressor entry.
The importance of the compressor pressure ratio is because overall engine performance is influenced by this parameter as it bears directly on thrust, fuel consumption, and engine efficiency. Engine weight, too, is directly related to pressure ratio. An increase in pressure ratio, for instance, may require the number of stages in the compressor to be increased, which will result in higher pressure and temperature levels within the gas generator. This in turn will necessitate a heavier engine overall because not only must the engine be re-designed to withstand the higher stress levels, but usually the combustion chamber and turbine as well.
The mass flow rate parameter denotes the airflow volume that the compressor is capable of processing within unit time, usually one second. Apart from the importance of this parameter for thermal cycle analysis, it also permits engine classification with respect to engine size to easily be made.
There are basically two types of compressors that are in use today.
- the centrifugal-flow compressor, and
- the axial-flow compressor.
Today centrifugal compressors are used only in small engines such as shaft engines for helicopters, auxiliary power units, turboprop engines, and some low-thrust engines for business aircraft. The majority of engines, however, employ the axial-flow compressor
The principal advantage of the axial compressor is its ability to deliver High Mass flow rates together with large pressure ratios at the same time-features which the centrifugal compressor, due to its method of compression, cannot provide.
The axial compressor is usually made up of a large number of individual parts of diverse function. Although the variety of engine makes available on the market differ according to the requirements to individual applications, distinct types of components are typical to any compressor. These are
- compressor front frame
- compressor casing with stator vanes
- rotor with rotor blades
- compressor rear frame.
The airflow, after being delivered to the compressor face by the air intake duct, first passes the front frame. This is a ring-shaped single-piece lightweight structure made of aluminium alloy or steel, usually cast and then machined. Characteristic to this component is an outer ring, an inner hub and 6 to 8 streamlined supporting struts.
The task of the compressor front frame is to accommodate the rotor front bearing and to transfer rotor forces to the outer casing by means of the supporting struts.
The supporting struts are hollow to accommodate, tubing to lubricate and ventilate the front bearing, and to provide space for electric cables where an electric starter is mounted forward of the shaft. The interior of the hollow struts are also used as a warm air passage for de-icing the struts themselves.
The compressor casing is a tube-like construction typically split length-wise to facilitate engine assembly and maintenance. The casing material is usually lightweight titanium forging, but stainless steel has also been used in the past. Modern high-performance engines employ alloy materials that were specifically developed to allow expansion of the case due to heating during engine operation.
A proportion of the compressed air is permanently bled, either through circumferentially arranged bleed manifolds or through hollow stator vanes. Bleed air is used for aircraft systems such as cabin pressurization and heating, wing leading edge de-icing, and for electronic systems temperature control. The engine itself also requires bleed air both for de-icing the front frame support struts and nose cowl, and for cooling the turbine frame and blades.
Guide vanes are used to impose a direction to the flow, and to convert rotor exit swirl velocity into a static pressure rise. The number of guide vanes within a compressor assembly may be substantial, in some cases several hundred.
The rotor is considered the most complex component of the compressor assembly. Energies of several ten thousands of horsepower may be processed in some compressors, in particular those of high bypass-engines. Such sever load condition require4 unique methods of rotor construction.
In its general design the rotor may be of the drum or disc type, or be a combination shaft and disc structure. In a disc-type rotor the rotor blades are mounted on individual discs, which are then separately secured to the rotor shaft, often divided by spacer rings. Individual construction varies with engine manufacturer, but the principle of transferring torque and axial loads at the same time is characteristic of any axial-flow rotor.
The shape of the rotor blades is comparable to a miniature wing featuring the typical aerofoil section. Unlike an aircraft wing, however, a rotor blade may be highly twisted from root to tip to obtain the optimum angle-of-attack to the flow everywhere along the blade length. The reason is that the root section travels much slower than the tip section and ‘views’ the flow from a different direction. The necessity for blade twist arises from the requirement for constant axial velocity being maintained across the flow path. The length of the blades decreased progressively downstream in the same proportion as the pressure increase.
The basic airflow function of the compressor rear frame is to guide and deliver the pressurized airstreams to the combustion section. Flow path design, therefore, reflects the type of combustor employed. When can-type combustion chambers were used, the flow path through the rear frame had to be equally apportioned to each combustor. The cross-section of the flow path progressively increases downstream to act as a diffuser, i.e. to reduce airstreams velocity and increase pressure.
With regard to engine thrust forces, the compressor rear frame is of great importance. In most cases it is here where the primary engine mounting is located and thrust forces are transmitted to the airframe.
The center of the compressor rear frame is designed to house the rearward bearing of the rotor, a ball bearing that absorbs the longitudinal thrust of the rotor. With high-thrust engines this bearing must withstand extreme loads.
The struts of the compressor rear frame, in addition to contributing structural strength to the compressor assembly, may also serve to facilitate lubrication and venting of the bearing as well as the supply of bleed air.
Axial compressors typically contain between 8 and 16 stages. A stage is the term given to a particular turbo-machinery unit in the compressor or the turbine. In the compressor, a stage consists of a rotor wheel carrying rotating blades, followed by a stator assembly carrying stationary blades or vanes.
The necessity for fuel to be burnt at the highest level of efficiency is fundamental in the aero gas turbine engine. Combustion efficiency directly affects the fuel load aircraft weight payload equation, and therefore the operation costs and range performance. Added to this are environmental problems calling for a reduction of dangerous emissions that result from combustion?
The development of combustion chambers is based essentially on experience with previous systems of similar design. In spite of a multitude of possible solutions for a particular combustion system, certain principals of design will be found in any combustion chamber.
The basic task of the combustion chamber is to provide a stream of hot gas that is able to release its energy to the turbine and nozzle sections of the engine. Following an increase in pressure through the compressor section, Heat is added to the airflow by the burning of a combustible gaseous mixture of vaporized fuel and highly compressed air. The combustion chambers and must be accomplished at a minimal loss of pressure.
The air mass flow when discharged from the compressor enters the combustion chamber at a velocity of around 150m/s (490ft/s)-far too high to sustain a flame for combustion. What is required in the first place is a slowing down of the aircraft. This is achieved in the forward section of the combustion chamber, which is formed as a diffuser; that is, the flow passage cross-section increases in the downstream direction. The result not only is a decrease in airflow velocity, but at the same time a further increase in pressure. Airflow velocity is now around 25m/s (80ft/s), still too high for orderly burning of the kerosene/air mixture. Flow velocity, therefore, must be further diminished down to a few meters per second. This is accomplished by means of a perforated disk that surrounds the fuel.
The second essential task of the combustion chamber is to provide the correct fuel/air mixture. The mass ratio of the two components that react in the combustion process, namely fuel mass injected per second and air mass forced each second into the combustion chamber, varies with the operation conditions of the aircraft and may range between ratios of 1:45 to1: 130. The fuel/air ratio for efficient combustion, however, is in the order of 1: 15, which means that only a fraction of the incoming air is required for the combustion process. The task of reducing flow velocity for the orderly burning of the fuel and apportioning the airflow to achieve complete combustion is accomplished in the forward section of the combustion chamber.
Apportioning the air for combustion is achieved by means of a short air duct, which has a number of drag-producing swirl vanes at the exit to reduce flow velocity. Airflow passing through the snout is only 20 per cent of the total mass of air entering the combustion chamber. By far the largest part is ducted around the internal flame tube, from where gradual admixing within the flame tube is made by means of various- size holes arranged behind the primary combustion zone.
Fuel is pumped into the injection nozzle at high pressure. The form of the injection nozzle ensures that the vaporized fuel is discharged as a spray cone, which provides intensive mixing with the air passing by. Fuel-burning takes place in a relatively small space within the flame tube, the primary combustion zone, where temperatures may be as high as 2,000K (3,600R). No flame tube material would be able to withstand such temperatures if the walls were not intensively cooled.
To this end a system of small holes and slots in the liner wall allows secondary cooling air to provide a protective shielding order to insulate the flame tube walls from the super-hot flames. The remaining part of the secondary air is ducted along the flame tube and gradually added to the hot gas. The combustion process must have ended before this to prevent incomplete combustion due to ‘low’ temperatures. Combustion is usually initiated by electrical spark ignition and then continues a self-sustaining process.
The can-type combustion chamber is found in early jet engines. First of single burners are arranged in parallel circumferentially around the engine axis. Each chamber is supplied with a stream of airflow by a separate air duct that connects upstream to the compressor outlet. Burners are linked by interconnections that enable the flame to spread to neighbouring combustors, thus igniting the fuel/air mixture there, whereas start-up ignition is made only at two combustors. The interconnections also act in equalizing the pressure among all burner cans to ensure identical operating conditions in all combustors and thereby prevent asymmetric turbine loading.
Because of the inefficient use of space in unfavourable fluid dynamic effects, the can-type burner is no longer used in aero-engines. Today, an annular-type combustor provides the most efficient use of space. The annular-type burner is a single concentric flame tube surrounding the spools. This results in a 25 percent reduction in weight as compared to the can-type burner. Being a single large combustion chamber, the process of combustion is more evenly achieved within the flame tube.
The primary task of the turbine in an aero-engine is to drive the compressor. Additionally, the turbine must drive the accessories. Basically turbine operation is no different from that of a compressor. A turbine absorbs energy from the gas flow to convert it into mechanical shaft power or torque. In aero engines, the axial-type turbine is exclusively used because of the higher mass flow rate it makes possible.
A turbine stage comprises two main elements consisting of:
a) a set of stationary nozzle guide vanes followed downstream by
b) a set of rotating blades. (This sequence is reversed in a compressor.)
In order to perform work, the hot gas discharged from the combustion chamber must be suitably processed. This is the task of nozzle guide vanes, and they have two principals’ functions. First, they must convert part of the energy of the hot gas into kinetic energy in order to make the flow fast enough when it impinges on the rotor blades. Second, the nozzle guide vanes must change the direction of the gas flow in a manner such that the circumferential forces engendered in the blades are maximized for the production of shaft power.
The required acceleration is accomplished by narrowing the passage between adjacent blades. As velocity increases, static pressure and temperature decrease. The degree of this energy conversion depends on the relationship of nozzle inlet to exit area, which is a direct function of the type of turbine blade used.
As no work is done by the hot gas in the nozzle guide vane section, the gas total energy will remain constant, if flow losses are neglected. It is only the state of part of the energy that is changing from the potential to the kinetic, i.e. heat and pressure energy are converted into gas velocity energy.
During the expansion of the gas in a turbine, energy contained in the gas is extracted and converted into mechanical energy, in the form of shaft power. The amount of energy absorbed by the turbine is only as much as required for driving the compressor and accessories such as the fuel pump, oil pump, electric generator. In engines used for jet propulsion a large proportion of gas energy is still available to be converted into engine thrust.
The task of the exhaust nozzle is to convert gas potential energy into kinetic energy necessary for the generation of thrust. This is accomplished solely by the geometrical shape of the nozzle, which is basically a tube of varying cross-section.
As most aircraft land at speeds around 250 km/h, long runway distances are required. Apart from high landing speed, aircraft weight and the limited capacity of mechanical wheel brakes all compound the slowing-down problem.
It appeared logical to use the energy in the propulsion system to slow the aircraft down. The result is the thrust reverser, which has become an integral part of the exhaust system.
The thrust reverser functions by obstructing the exhaust by blocker doors, which can be turned into the flow. Turning the exhaust flow to a forward direction results in a forward thrust, which acts as a brake. The design of a particular thrust reverser depends on the engine with which it is used. In all cases, the only engine which will use a thrust reverser will be of the turbofan type.
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