# Jet engine (المحرك النفاث) شرح بالصور مدعوم بروابط متداخلة



## حسن هادي (21 أغسطس 2007)

سنتطرق عن المحرك النفاث في هذا الموضوع استكمالا لما بدانا به من مواضيع التوربينات ومكائن الاحتراق الداخلي وغيرها وكما اسلفنا بالذكر بان تداخل هذه المواضيع من الناحية العلمية والعملية يدفعنا الى ربط بعضها بالاخر لنسهل على اخوتنا عملية البحث ومن الله التوفيق /اخوكم حسن 
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*Jet engine*

*From Wikipedia:, the free encyclopedia موسوعة رائعة: *


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A Pratt & Whitney F100 turbofan engine for the F-15 Eagle and the F-16 Falcon is tested at Robins Air Force Base, Georgia, USA. The tunnel behind the engine muffles noise and allows exhaust to escape.


A *jet engine* is an engine that discharges a fast moving jet of fluid to generate thrust in accordance with Newton's third law of motion. This broad definition of jet engines includes turbojets, turbofans, rockets, ramjets, pulse jets and pump-jets, but in common usage, the term generally refers to a gas turbine Brayton cycle engine, an engine with a rotary compressor powered by a turbine, with the leftover power providing thrust. Jet engines are so familiar to the modern world that gas turbines are sometimes mistakenly referred to as a particular application of a jet engine, rather than the other way around. Most jet engines are internal combustion engines but non combusting forms exist.


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## حسن هادي (21 أغسطس 2007)

Jet engines can be dated back to the first century AD, when Hero of Alexandria invented the aeolipile. This used steam power directed through two jet nozzles so as to cause a sphere to spin rapidly on its axis. So far as is known, it was little used for supplying mechanical power, and the potential practical applications of Hero's invention of the jet engine were not recognized. It was simply considered a curiosity.
Jet propulsion only literally and figuratively took off with the invention of the rocket by the Chinese in the 11th century. Rocket exhaust was initially used in a modest way for fireworks but gradually progressed to propel formidable weaponry; and there the technology stalled for hundreds of years.
The problem was that rockets are simply too inefficient at low speeds to be useful for general aviation. Instead, by the 1930s, the piston engine in its many different forms (rotary and static radial, aircooled and liquid-cooled inline) was the only type of powerplant available to aircraft designers. This was acceptable as long as only low performance aircraft were required, and indeed all that were available.
However, engineers were beginning to realize that the piston engine was self-limiting in terms of the maximum performance which could be attained; the limit was essentially one of propeller efficiency.[1] This seemed to peak as blade tips approached the speed of sound. If engine, and thus aircraft, performance were ever to increase beyond such a barrier, a way would have to be found to radically improve the design of the piston engine, or a wholly new type of powerplant would have to be developed. This was the motivation behind the development of the gas turbine engine, commonly called a "jet" engine, which would become almost as revolutionary to aviation as the Wright brothers' first flight.
The earliest attempts at jet engines were hybrid designs in which an external power source first compressed air, which was then mixed with fuel and burned for jet thrust. In one such system, called a thermojet by Secondo Campini but more commonly, motorjet, the air was compressed by a fan driven by a conventional piston engine. Examples of this type of design were Henri Coandă's Coandă-1910 aircraft, and the much later Campini Caproni CC.2, and the Japanese Tsu-11 engine intended to power Ohka kamikaze planes towards the end of World War II. None were entirely successful and the CC.2 ended up being slower than the same design with a traditional engine and propeller combination.


 


simulation of the Jet Engine Airflow


The key to a practical jet engine was the gas turbine, used to extract energy from the engine itself to drive the compressor. The gas turbine was not an idea developed in the 1930s: the patent for a stationary turbine was granted to John Barber in England in 1791. The first gas turbine to successfully run self-sustaining was built in 1903 by Norwegian engineer Ægidius Elling. The first patents for jet _propulsion_ were issued in 1917. Limitations in design and practical engineering and metallurgy prevented such engines reaching manufacture. The main problems were safety, reliability, weight and, especially, sustained operation.


 


The W2/700 engine flew in the Gloster E.28/39, the first British aircraft to fly with a turbojet engine, and the Gloster Meteor.


In 1929, Aircraft apprentice Frank Whittle formally submitted his ideas for a turbo-jet to his superiors. On 16 January 1930 in England, Whittle submitted his first patent (granted in 1932). The patent showed a two-stage axial compressor feeding a single-sided centrifugal compressor. Whittle would later concentrate on the simpler centrifugal compressor only, for a variety of practical reasons. Whittle had his first engine running in April 1937. It was liquid-fuelled, and included a self-contained fuel pump. Whittle's team experienced near-panic when the engine would not stop, even after the fuel was switched off. It turned out that fuel had leaked into the engine and accumulated in pools. So the engine would not stop until all the leaked fuel had burned off. Whittle was unable to interest the government in his invention, and development continued at a slow pace.
In 1935 Hans von Ohain started work on a similar design in Germany, unaware of Whittle's work. His first engine was strictly experimental and could only run under external power, but he was able to demonstrate the basic concept. Ohain was then introduced to Ernst Heinkel, one of the larger aircraft industrialists of the day, who immediately saw the promise of the design. Heinkel had recently purchased the Hirth engine company, and Ohain and his master machinist Max Hahn were set up there as a new division of the Hirth company. They had their first HeS 1 engine running by September 1937. Unlike Whittle's design, Ohain used hydrogen as fuel, supplied under external pressure. Their subsequent designs culminated in the gasoline-fuelled HeS 3 of 1,100 lbf (5 kN), which was fitted to Heinkel's simple and compact He 178 airframe and flown by Erich Warsitz in the early morning of August 27, 1939, from Marienehe aerodrome, an impressively short time for development. The He 178 was the *world's first jet plane.*
Meanwhile, Whittle's engine was starting to look useful, and his *Power Jets Ltd.* started receiving Air Ministry money. In 1941 a flyable version of the engine called the *W.1*, capable of 1000 lbf (4 kN) of thrust, was fitted to the Gloster E28/39 airframe specially built for it, and first flew on May 15, 1941 at RAF Cranwell.


 


A picture of an early centrifugal engine (the _DH Goblin II_) sectioned to show its internal components


One problem with both of these early designs, which are called *centrifugal-flow* engines, was that the compressor worked by "throwing" (accelerating) air outward from the central intake to the outer periphery of the engine, where the air was then compressed by a divergent duct setup, converting its velocity into pressure. An advantage of this design was that it was already well understood, having been implemented in centrifugal superchargers, then in widespread use on piston engines. However, given the early technological limitations on the shaft speed of the engine, the compressor needed to have a very large diameter to produce the power required. This meant that the engines had a large frontal area, which made it less useful as an aircraft powerplant due to drag. A further disadvantage was that the air flow had to be "bent" to flow rearwards through the combustion section and to the turbine and tailpipe, adding complexity and lowering efficiency. Nevertheless, Whittle's engines had the major advantages of light weight, simplicity and reliability, and development rapidly progressed to practical airworthy designs.
Austrian Anselm Franz of Junkers' engine division (_Junkers Motoren_ or *Jumo*) addressed these problems with the introduction of the axial-flow compressor. Essentially, this is a turbine in reverse. Air coming in the front of the engine is blown towards the rear of the engine by a fan stage (convergent ducts), where it is crushed against a set of non-rotating blades called _stators_ (divergent ducts). The process is nowhere near as powerful as the centrifugal compressor, so a number of these pairs of fans and stators are placed in series to get the needed compression. Even with all the added complexity, the resulting engine is much smaller in diameter and thus, more aerodynamic. Jumo was assigned the next engine number in the RLM numbering sequence, 4, and the result was the Jumo 004 engine. After many lesser technical difficulties were solved, mass production of this engine started in 1944 as a powerplant for the world's first jet-fighter aircraft, the Messerschmitt Me 262 (and later the worlds first jet-bomber aircraft, the Arado Ar 234). A variety of reasons conspired to delay the engine's availability, this delay caused the fighter too arrive too late to decisively impact Germany's position in World War II. Nonetheless, it will be remembered as the first use of jet engines in service. Following the end of the war the German jet aircraft and jet engines were extensively studied by the victorious allies and contributed to work on early Soviet and US jet fighters. The legacy of the axial-flow engine is seen in the fact that practically all jet engines on fixed wing aircraft have had some inspiration from this design.


 


A cutaway of the Junkers Jumo 004 engine.


Centrifugal-flow engines have improved since their introduction. With improvements in bearing technology the shaft speed of the engine was increased, greatly reducing the diameter of the centrifugal compressor. The short engine length remains an advantage of this design, particularly for use in helicopters where overall size is more important than frontal area. Also, its engine components are robust; axial-flow compressors are more liable to foreign object damage.
Although German designs were more advanced aerodynamically, the combination of simplicity and advanced British metallurgy meant that Whittle-derived designs were far more reliable than their German counterparts. British engines also were licensed widely in the US (see Tizard Mission), and were sent to the USSR in a technology exchange, with the Nene going on to power the famous MiG-15. American and Soviet designs, independent axial-flow types for the most part, would not come fully into their own until the 1960s, although the General Electric J47 provided excellent service in the F-86 Sabre in the 1950s.
By the 1950s the jet engine was almost universal in combat aircraft, with the exception of cargo, liaison and other specialty types. By this point some of the British designs were already cleared for civilian use, and had appeared on early models like the the deHavilland Comet and Canadair Jetliner. By the 1960s all large civilian aircraft were also jet powered, leaving the piston engine in niche roles here as well. Relentless improvements in the turboprop has since pushed the piston engine out of the mainstream entirely, leaving it serving only the smallest general aviation designs, and some use in drone aircraft. The ascension of the jet engine to almost universal use in aircraft use took well under twenty years.


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## حسن هادي (21 أغسطس 2007)

*Type comparison*



 


Comparative suitability for (left to right) turboshaft, low bypass and turbojet to fly at 10 km attitude in various speeds. Horizontal axis - speed, m/s. Vertical axis displays engine efficiency.




 


Specific impulse as a function of speed of different Jet types. Although efficiency plummets with speed, greater distances are covered, it turns out that efficiency per unit distance (per km or mile) is roughly independent of speed for Jet engines as a group; however airframes become inefficient at supersonic speeds




 


Dependence of the energy efficiency (η) from the exhaust speed/airplane speed ratio (c/v) for airbreathing jets




 


Dependence of the energy efficiency (η) upon the vehicle speed/exhaust speed ratio (v/c) for rocket engines


The motion impulse of the engine is equal to the air mass multiplied by the speed at which the engine emits this mass:
I = m c where m is the air mass per second and c is the exhaust speed. In other words, the plane will fly faster if the engine emits the air mass with a higher speed or if it emits more air per second with the same speed. However, when the plane flies with certain velocity v, the air moves towards it, creating the opposing ram drag at the air intake:
m v Most types of jet engine have an air intake, which provides the bulk of the gas exiting the exhaust. Conventional rocket motors, however, do not have an air intake, the oxidizer and fuel both being carried within the airframe. Therefore, rocket motors do not have ram drag; the gross thrust of the nozzle is the net thrust of the engine. Consequently, the thrust characteristics of a rocket motor are completely different from that of an air breathing jet engine.
The air breathing engine is only useful if the velocity of the gas from the engine, c, is greater than the airplane velocity, v. The net engine thrust is the same as if the gas were emitted with the velocity c-v. So the thrust is actually equal to
S = m (c-v) The turboprop has a wide rotating fan that takes and accelerates the large mass of air but only till the limited speed of any propeller driven airplane. When the plane speed exceeds this limit, propellers no longer provide any thrust (c-v < 0).
The turbojets and other similar engines accelerate much smaller mass of the air and burned fuel, but they emit it at the much higher speeds possible with a de Laval nozzle. This is why they are suitable for supersonic and higher speeds.
From the other side, the _propulsive efficiency_ (essentially energy efficiency) is highest when the engine emits an exhaust jet at a speed that is the same as the airplane velocity. The exact formula, given in the literature,[3] is




The low bypass turbofans have the mixed exhaust of the two air flows, running at different speeds (c1 and c2). The thrust of such engine is
S = m1 (c1 - v) + m2 (c2 - v) where m1 and m2 are the air masses, being blown from the both exhausts. Such engines are effective at lower speeds, than the pure jets, but at higher speeds than the turboshafts and propellers in general. For instance, at the 10 km attitude, turboshafts are most effective at about 0.4 mach, low bypass turbofans become more effective at about 0.75 mach and true jets become more effective as mixed exaust engines when the speed approaches 1 mach - the speed of sound.
Rocket engines are best suited for high speeds and altitudes. At any given throttle, the thrust and efficiency of a rocket motor improves slightly with increasing altitude (because the back-pressure falls thus increasing net thrust at the nozzle exit plane), whereas with a turbojet (or turbofan) the falling density of the air entering the intake (and the hot gases leaving the nozzle) causes the net thrust to decrease with increasing altitude. Rocket engines are more efficient than even scramjets above roughly Mach 15.[4]


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## حسن هادي (21 أغسطس 2007)

Turbojet engines


 


A turbojet engine, in its simplest form is simply a gas turbine with a nozzle attached


_Main article: Turbojet_
A turbojet engine is a type of internal combustion engine often used to propel aircraft. Air is drawn into the rotating compressor via the intake and is compressed, through successive stages, 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 both the gas temperature and pressure at exit from the turbine, 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.
Under normal circumstances, the pumping action of the compressor prevents any backflow, thus facilitating the continuous-flow process of the engine. Indeed, the entire process is similar to a four-stroke cycle, but with induction, compression, ignition, expansion and exhaust taking place simultaneously, but in different sections of the engine. The efficiency of a jet engine is strongly dependent upon the overall pressure ratio (combustor entry pressure/intake delivery pressure) and the turbine inlet temperature of the cycle.
It is also perhaps instructive to compare turbojet engines with propeller engines. Turbojet engines take a relatively small mass of air and accelerate it by a large amount, whereas a propeller takes a large mass of air and accelerates it by a small amount. The high-speed exhaust of a turbojet engine makes it efficient at high speeds (especially supersonic speeds) and high altitudes. On slower aircraft and those required to fly short stages, a gas turbine-powered propeller engine, commonly known as a turboprop, is more common and much more efficient. Very small aircraft generally use conventional piston engines to drive a propeller but small turboprops are getting smaller as engineering technology improves.
The turbojet described above is a single-spool design, in which a single shaft connects the turbine to the compressor. Higher overall pressure ratio designs often have two concentric shafts, to improve compressor stability during engine throttle movements. The outer high pressure (HP) shaft connects the HP compressor to the HP turbine. This HP Spool, with the combustor, forms the core or gas generator of the engine. The inner shaft connects the low pressure (LP) compressor to the LP Turbine to create the LP Spool. Both spools are free to operate at their optimum shaft speed. (Concorde used this type).


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## حسن هادي (21 أغسطس 2007)

دورة برايتون مع المخططات/
 تحياتي اخوكم *


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## حسن هادي (21 أغسطس 2007)

Turbofan engines
_Main article: Turbofan_
Most modern jet engines are actually turbofans, where the low pressure compressor acts as a fan, supplying supercharged air to not only the engine core, but to a bypass duct. The bypass airflow either passes to a separate 'cold nozzle' or mixes with low pressure turbine exhaust gases, before expanding through a 'mixed flow nozzle'.
Turbofans are used for airliners because they give an exhaust speed that is better matched to subsonic airliner's flight speed, conventional turbojet engines generate an exhaust that ends up travelling very fast backwards, and this wastes energy. By emitting the exhaust so that it ends up travelling more slowly, better fuel consumption is achieved. In addition, the lower exhaust speed gives much lower noise.
In the 1960s there was little difference between civil and military jet engines, apart from the use of afterburning in some (supersonic) applications. Civil turbofans today have a low specific thrust (net thrust divided by airflow) to keep jet noise to a minimum and to improve fuel efficiency. Consequently the bypass ratio (bypass flow divided by core flow) is relatively high (ratios from 4:1 up to 8:1 are common). Only a single fan stage is required, because a low specific thrust implies a low fan pressure ratio.
Today's military turbofans, however, have a relatively high specific thrust, to maximize the thrust for a given frontal area, jet noise being of less concern in military uses relative to civil uses. Multistage fans are normally needed to reach the relatively high fan pressure ratio needed for high specific thrust. Although high turbine inlet temperatures are often employed, the bypass ratio tends to be low, usually significantly less than 2.0.
An approximate equation for calculating the net thrust of a jet engine, be it a turbojet or a mixed turbofan, is:




where:



intake mass flow rate



fully expanded jet velocity (in the exhaust plume)



aircraft flight velocity
While the



term represents the gross thrust of the nozzle, the



term represents the ram drag of the intake.


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## حسن هادي (21 أغسطس 2007)




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## سنان عبد الغفار (21 أغسطس 2007)

كلمة شكرا لا توفي ماتقدمه ايها العضو المميز حسن هادي اسأل الله ان يرزقك ويهنأك بحياتك لان ثلاثة يسأل عنها الله يوم القيامة احدها (علماً ينتفع به)
ونرجو منك المزيك وتفاصل اكثر حول موضوع التوربينات النفاثة


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## حسن هادي (21 أغسطس 2007)

Major components


 


Basic components of a jet engine (Axial flow design)


The major components of a jet engine are similar across the major different types of engines, although not all engine types have all components. The major parts include:

*Cold Section:*
*Air intake (Inlet)* — The standard reference frame for a jet engine is the aircraft itself. For subsonic aircraft, the air intake to a jet engine presents no special difficulties, and consists essentially of an opening which is designed to minimise drag, as with any other aircraft component. However, the air reaching the compressor of a normal jet engine must be travelling below the speed of sound, even for supersonic aircraft, to sustain the flow mechanics of the compressor and turbine blades. At supersonic flight speeds, shockwaves form in the intake system and reduce the recovered pressure at inlet to the compressor. So some supersonic intakes use devices, such as a cone or ramp, to increase pressure recovery, by making more efficient use of the shock wave system.
*Compressor* or *Fan* — The compressor is made up of stages. Each stage consists of vanes which rotate, and stators which remain stationary. As air is drawn deeper through the compressor, its heat and pressure increases. Energy is derived from the *turbine* (see below), passed along the *shaft*.

*Common:*
*Shaft* — The shaft connects the *turbine* to the *compressor*, and runs most of the length of the engine. There may be as many as three concentric shafts, rotating at independent speeds, with as many sets of turbines and compressors. Other services, like a bleed of cool air, may also run down the shaft.

*Hot section:*
*Combustor* or *Can* or *Flameholders* or *Combustion Chamber* — This is a chamber where fuel is continuously burned in the compressed air.
*Turbine* — The turbine acts like a windmill, gaining energy from the hot gases leaving the *combustor*. This energy is used to drive the *compressor* (or props, or bypass fans) via the *shaft*, or even (for a gas turbine-powered helicopter) converted entirely to rotational energy for use elsewhere. Relatively cool air, bled from the compressor, may be used to cool the turbine blades and vanes, to prevent them from melting.
*Afterburner* or *reheat* (chiefly UK) — (mainly military) Produces extra thrust by burning extra fuel, usually inefficiently, to significantly raise Nozzle Entry Temperature at the *exhaust*. Owing to a larger volume flow (i.e. lower density) at exit from the afterburner, an increased nozzle flow area is required, to maintain satisfactory engine matching, when the afterburner is alight.
*Exhaust* or *Nozzle* — Hot gases leaving the engine exhaust to atmospheric pressure via a nozzle, the objective being to produce a high velocity jet. In most cases, the nozzle is convergent and of fixed flow area.
*Supersonic nozzle* — If the Nozzle Pressure Ratio (Nozzle Entry Pressure/Ambient Pressure) is very high, to maximize thrust it may be worthwhile, despite the additional weight, to fit a convergent-divergent (de Laval) nozzle. As the name suggests, initially this type of nozzle is convergent, but beyond the throat (smallest flow area), the flow area starts to increase to form the divergent portion. The expansion to atmospheric pressure and supersonic gas velocity continues downstream of the throat, whereas in a convergent nozzle the expansion beyond sonic velocity occurs externally, in the exhaust plume. The former process is more efficient than the latter.

The various components named above have constraints on how they are put together to generate the most efficiency or performance. The performance and efficiency of an engine can never be taken in isolation; for example fuel/distance efficiency of a supersonic jet engine maximises at about mach 2, whereas the drag for the vehicle carrying it is increasing as a square law and has much extra drag in the transonic region. The highest fuel efficiency for the overall vehicle is thus typically at Mach ~0.85.
For the engine optimisation for its intended use, important here is air intake design, overall size, number of compressor stages (sets of blades), fuel type, number of exhaust stages, metallurgy of components, amount of bypass air used, where the bypass air is introduced, and many other factors. For instance, let us consider design of the air intake.


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## حسن هادي (21 أغسطس 2007)

سنان عبد الغفار قال:


> كلمة شكرا لا توفي ماتقدمه ايها العضو المميز حسن هادي اسأل الله ان يرزقك ويهنأك بحياتك لان ثلاثة يسأل عنها الله يوم القيامة احدها (علماً ينتفع به)
> ونرجو منك المزيك وتفاصل اكثر حول موضوع التوربينات النفاثة


 
حياك الله اخي العزيز سنان عبد الغفار وتقبل مني كل المودة والتقدير


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## حسن هادي (21 أغسطس 2007)

No higher resolution available.
Axial_compressor.gif (358 × 295 pixel, file size: 172 KB, MIME type: image/gif)
الاخوة الاعزاء الروابط فعالة وتقبلوا تحياتي*


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## حسن هادي (21 أغسطس 2007)

* Air intakes*

_See also: Inlet cone_

*Subsonic inlets*



 


Pitot intake operating modes


Pitot intakes are the dominant type for subsonic applications. A subsonic pitot inlet is little more than a tube with an aerodynamic fairing around it.
At zero airspeed (i.e., rest), air approaches the intake from a multitude of directions: from directly ahead, radially, or even from behind the plane of the intake lip.
At low airspeeds, the streamtube approaching the lip is larger in cross-section than the lip flow area, whereas at the intake design flight Mach number the two flow areas are equal. At high flight speeds the streamtube is smaller, with excess air spilling over the lip.
Beginning around 0.85 Mach, shock waves can occur as the air accelerates through the intake throat.
Careful radiusing of the lip region is required to optimize intake pressure recovery (and distortion) throughout the flight envelope.

*[edit] Supersonic inlets*

Supersonic intakes exploit shock waves to decelerate the airflow to a subsonic condition at compressor entry.
There are basically two forms of shock waves:
1) Normal shock waves lie perpendicular to the direction of the flow. These form sharp fronts and shock the flow to subsonic speeds. Microscopically the air molecules smash into the subsonic crowd of molecules like alpha rays. Normal shock waves tend to cause a large drop in stagnation pressure. Basically, the higher the supersonic entry Mach number to a normal shock wave, the lower the subsonic exit Mach number and the stronger the shock (i.e. the greater the loss in stagnation pressure across the shock wave).
2) Conical (3-dimensional) and oblique shock waves (2D) are angled rearwards, like the bow wave on a ship or boat, and radiate from a flow disturbance such as a cone or a ramp. For a given inlet Mach number, they are weaker than the equivalent normal shock wave and, although the flow slows down, it remains supersonic throughout. Conical and oblique shock waves turn the flow, which continues in the new direction, until another flow disturbance is encountered downstream.
Note: Comments made regarding 3 dimensional conical shock waves, generally also apply to 2D oblique shock waves.
A sharp-lipped version of the pitot intake, described above for subsonic applications, performs quite well at moderate supersonic flight speeds. A detached normal shock wave forms just ahead of the intake lip and 'shocks' the flow down to a subsonic velocity. However, as flight speed increases, the shock wave becomes stronger, causing a larger percentage decrease in stagnation pressure (i.e. poorer pressure recovery). An early US supersonic fighter, the F-100 Super Sabre, used such an intake.


 


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


More advanced supersonic intakes, excluding pitots:
a) exploit a combination of conical shock wave/s and a normal shock wave to improve pressure recovery at high supersonic flight speeds. Conical shock wave/s are used to reduce the supersonic Mach number at entry to the normal shock wave, thereby reducing the resultant overall shock losses.
b) have a design shock-on-lip flight Mach number, where the conical/oblique shock wave/s intercept the cowl lip, thus enabling the streamtube capture area to equal the intake lip area. 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. Consequently, the intake capture area is less than the intake lip area, which reduces the intake airflow. Depending on the airflow characteristics of the engine, it may be desirable to lower the ramp angle or move the cone rearwards to refocus the shockwaves onto the cowl lip to maximise intake airflow.
c) are designed to have a normal shock in the ducting downstream of intake lip, so that the flow at compressor/fan entry is always subsonic. However, if the engine is throttled back, there is a reduction in the corrected airflow of the LP compressor/fan, but (at supersonic conditions) the corrected airflow at the intake lip remains constant, because it is determined by the flight Mach number and intake incidence/yaw. 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. This weakens the shockwave, improving the overall intake pressure recovery. So, the absolute airflow stays constant, whilst the corrected airflow at compressor entry falls (because of a higher entry pressure). Excess intake airflow may also be dumped overboard or into the exhaust system, to prevent the conical/oblique shock waves being disturbed by the normal shock being forced too far forward by engine throttling.
Many second generation supersonic fighter aircraft featured an inlet cone, which was used to form the conical shock wave. This type of inlet cone is clearly seen at the very front of the English Electric Lightning and MiG-21 aircraft, for example.
The same approach can be used for air intakes mounted at the side of the fuselage, where a half cone serves the same purpose with a semicircular air intake, as seen on the F-104 Starfighter and BAC TSR-2.
Some intakes are biconic; that is they feature two conical surfaces: the first cone is supplemented by a second, less oblique, conical surface, which generates an extra conical shockwave, radiating from the junction between the two cones. A biconic intake is usually more efficient than the equivalent conical intake, because the entry Mach number to the normal shock is reduced by the presence of the second conical shock wave.
A very sophisticated conical intake was featured on the SR-71's Pratt & Whitney J58s that could move a conical spike fore and aft within the engine nacelle, preventing the shockwave formed on the spike from entering the engine and stalling the engine, while keeping it close enough to give good compression. Movable cones are uncommon.
A more sophisticated design than cones is to angle the intake so that one of its edges forms a ramp. An oblique shockwave will form at the start of the ramp. The Century Series of US jets featured several variants of this approach, usually with the ramp at the outer vertical edge of the intake, which was then angled back inward towards the fuselage. Typical examples include the Republic F-105 Thunderchief and F-4 Phantom.


 


Concorde intake operating modes


Later this evolved so that the ramp was at the top horizontal edge rather than the outer vertical edge, with a pronounced angle downwards and rearwards. This design simplified the construction of intakes and allowed use of variable ramps to control airflow into the engine. Most designs since the early 1960s now feature this style of intake, for example the F-14 Tomcat, Panavia Tornado and Concorde.
From another point of view, like in a supersonic nozzle the corrected (or non-dimensional) flow has to be the same at the intake lip, at the intake throat and at the turbine. One of this three can be fixed. For inlets the throat is made variable and some air is bypassed around the turbine and directly fed into the afterburner. Unlike in a nozzle the inlet is either unstable or inefficient, because a normal shock wave in the throat will suddenly move to the lip, thereby increasing the pressure at the lip, leading to drag and reducing the pressure recovery, leading to turbine surge and the loss of one SR-71.


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## حسن هادي (21 أغسطس 2007)

نرفق لكم بعض الاضافات للموضوع لاستيفاء الفائدة وتحياتنا لكل الاعضاء**
*******************************************************************************
Compressors
Axial compressors rely on spinning blades that have aerofoil sections, similar to aeroplane wings. As with aeroplane wings in some conditions the blades can stall. If this happens, the airflow around the stalled compressor can reverse direction violently. Each design of a compressor has an associated operating map of airflow versus rotational speed for characteristics peculiar to that type (see compressor map).
At a given throttle condition, the compressor operates somewhere along the steady state running line. Unfortunately, this operating line is displaced during transients. Many compressors are fitted with anti-stall systems in the form of bleed bands or variable geometry stators to decrease the likelihood of surge. Another method is to split the compressor into two or more units, operating on separate concentric shafts.
Another design consideration is the average stage loading. This can be kept at a sensible level either by increasing the number of compression stages (more weight/cost) or the mean blade speed (more blade/disc stress).
Although large flow compressors are usually all-axial, the rear stages on smaller units are too small to be robust. Consequently, these stages are often replaced by a single centrifugal unit. Very small flow compressors often employ two centrifugal compressors, connected in series. Although in isolation centrifugal compressors are capable of running at quite high pressure ratios (e.g. 10:1), impeller stress considerations (i.e. T3, NH implications) limit the pressure ratio that can be employed in high overall pressure ratio engine cycles.
Increasing overall pressure ratio implies raising the high pressure compressor exit temperature (i.e. T3). 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.


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## حسن هادي (21 أغسطس 2007)

نرجوا ان يكون الموضوع مفيدا *علما ان جميع الروابط عاملة ضمن موسوعة ويكيبيديا مع التقدير

:6: :6: :6: :6:


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## شكرى محمد نورى (21 أغسطس 2007)

الأخ حسن هادي .

تحية طيبة .

ماتطرحه من مواضيع نادرة وممتازة وفي غاية من الروعة .

البغدادي .


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## حسن هادي (21 أغسطس 2007)

Combustors
Great care must be taken to keep the flame burning in a moderately fast moving airstream, at all throttle conditions, as efficiently as possible. Since the turbine cannot withstand stoichiometric temperatures, resulting from the optimum combustion process, some of the compressor air is used to quench the exit temperature of the combustor to an acceptable level. Air used for combustion is considered to be primary airflow, while excess air used for cooling is called secondary airflow. Combustor configurations include can, annular, and can-annular.

*[edit] Turbines*



 


Turbine Stage GE J79


Because a turbine expands from high to low pressure, there is no such thing as turbine surge or stall. 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. The blades have more curvature and the gas stream velocities are higher.
Designers must, however, prevent the turbine blades and vanes from melting in a very high temperature and stress environment. Consequently bleed air extracted from the compression system is often used to cool the turbine blades/vanes internally. Other solutions are improved materials and/or special insulating coatings. The discs must be specially shaped to withstand the huge stresses imposed by the rotating blades. They take the form of impulse, reaction, or combination impulse-reaction shapes. Improved materials help to keep disc weight down.

*[edit] Turbopumps*

_Main article: Turbopump_
Turbopumps are centrifugal pumps which are spun by gas turbines and are used to raise the propellant pressure above the pressure in the combustion chamber so that it can be injected and burnt. Turbopumps are very commonly used with rockets, but ramjets and turbojets also have been known to use them


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## حسن هادي (21 أغسطس 2007)

الاخوة الاعضاء نعيد الاشارة الى كون موضوع المحرك النفاث ذو ارتباط علمي وعملي مع مواضيع التوربينات ومكائن الاحتراق الداخلي وان كان الترابط نسبيا مما حدا بنا نشير الى ذلك رغبة في نشر الفئدة من خلال ما يتيسر لنا من مواقع النت او الملفات التي بحوزتنا *ولا بد ان اشير الى موسوعة الويكبيديا الحرة اذ ان موضوعاتها مميزة جدا وذات تشعبات كثيرة *تمنياتي للجميع بالتوفيق 
*****************************************************************************
نكمل لكم لتسهيل عملية البحث 
Nozzles


 


Afterburner GE J79


The primary object of a nozzle is to expand the exhaust stream to atmospheric pressure, thereby producing a high velocity jet, relative to the vehicle. If the fully expanded jet has a higher impulse than the moving aircraft, there will be a forward thrust on the airframe.
Simple convergent nozzles are used on many jet engines. If the nozzle pressure ratio is above the critical value (about 1.8:1) a convergent nozzle will choke, resulting in some of the expansion to atmospheric pressure taking place downstream of the throat (i.e. smallest flow area), in the jet wake. Although much of the gross thrust produced will still be from the jet momentum, additional (pressure) thrust will come from the imbalance between the throat static pressure and atmospheric pressure.
Many military combat engines incorporate an afterburner (or reheat) in the engine exhaust system. When the system is lit, the nozzle throat area must be increased, to accommodate the extra exhaust volume flow, so that the turbomachinery is unaware that the afterburner is lit. A variable throat area is achieved by moving a series of overlapping petals, which approximate the circular nozzle cross-section.
At high nozzle pressure ratios, the exit pressure is often above ambient and much of the expansion will take place downstream of a convergent nozzle, which is inefficient. Consequently, some jet engines (notably rockets) incorporate a convergent-divergent nozzle, to allow most of the expansion to take place against the inside of a nozzle to maximise thrust. However, unlike the fixed con-di nozzle used on a conventional rocket motor, when such a device is used on a turbojet engine it has to be a complex variable geometry device, to cope with the wide variation in nozzle pressure ratio encountered in flight and engine throttling. This further increases the weight and cost of such an installation.


 


Variable Exhaust Nozzle, on the GE F404-400 low-bypass turbofan installed on a Boeing F-18


The simpler of the two is the *ejector nozzle*, which creates an effective nozzle through a secondary airflow and spring-loaded petals. At subsonic speeds, the airflow constricts the exhaust to a convergent shape. As the aircraft speeds up, the two nozzles dilate, which allows the exhaust to form a convergent-divergent shape, speeding the exhaust gasses past Mach 1. More complex engines can actually use a tertiary airflow to reduce exit area at very low speeds. Advantages of the ejector nozzle are relative simplicity and reliability. Disadvantages are average performance (compared to the other nozzle type) and relatively high drag due to the secondary airflow. Notable aircraft to have utilized this type of nozzle include the SR-71, Concorde, F-111, and Saab Viggen
For higher performance, it is necessary to use an *iris nozzle*. This type uses overlapping, hydraulically adjustable "petals". Although more complex than the ejector nozzle, it has significantly higher performance and smoother airflow. As such, it is employed primarily on high-performance fighters such as the F-14, F-15, F-16, though is also used in high-speed bombers such as the B-1B. Some modern iris nozzle additionally have the ability to change the angle of the thrust (see thrust vectoring).


 


Iris vectored thrust nozzle


Rocket motors also employ convergent-divergent nozzles, but these are usually of fixed geometry, to minimize weight. Because of the much higher nozzle pressure ratios experienced, rocket motor con-di nozzles have a much greater area ratio (exit/throat) than those fitted to jet engines.
At the other extreme, some high bypass ratio civil turbofans use an extremely low area ratio (less than 1.01 area ratio), convergent-divergent, nozzle on the bypass (or mixed exhaust) stream, to control the fan working line. The nozzle acts as if it has variable geometry. At low flight speeds the nozzle is unchoked (less than a Mach number of unity), so the exhaust gas speeds up as it approaches the throat and then slows down slightly as it reaches the divergent section. Consequently, the nozzle exit area controls the fan match and, being larger than the throat, pulls the fan working line slightly away from surge. At higher flight speeds, the ram rise in the intake increases nozzle pressure ratio to the point where the throat becomes choked (M=1.0). Under these circumstances, the throat area dictates the fan match and being smaller than the exit pushes the fan working line slightly towards surge. This is not a problem, since fan surge margin is much better at high flight speeds.


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## حسن هادي (21 أغسطس 2007)

شكرى محمد نورى قال:


> الأخ حسن هادي .
> 
> تحية طيبة .
> 
> ...


الاخ العزيز المهندس والمشرف شكري البغدادي اعتز بهذا الاطراء ووفقنا الله لما فيه الصلاح*


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## الطموني (22 أغسطس 2007)

و الله مشعارف شو اقلك بس كثر الله من امثالك و اعمالك وبارك لك في الدنيا والاخرة


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## ouadah (22 يوليو 2009)

mmmmmmmmmeeeeeeeeeeerrrrrrrrrrccccciiiiiiiiiii


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