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... tion. A proportion of the compressed air is permanently bled, either through circumferential ly 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 require 4 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 aerosol 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 airstream's 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 airstream's 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 150 m / s (490 ft / 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 25 m / s (80 ft / 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 to 1: 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, 000 K (3, 600 R). 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 circumferential ly 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 neighboring combustor's, thus igniting the fuel / air mixture there, whereas start-up ignition is made only at two combustor's. The interconnections also act in equalizing the pressure among all burner cans to ensure identical operating conditions in all combustor's and thereby prevent asymmetric turbine loading.
Because of the inefficient use of space in unfavorable 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 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. REFERCES. Hunecke, Klaus.
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