Turbina a gas

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Prof. Gilberto MateranoGas Turbine CombustionCHAPTER I: IntroductionEarly combustor developmentThe world war ii

Gloster Meteor fighterit was generally accepted that the piston engine had reached its limit as a propulsion system for high-speed flight and the gas turbine was firmly established as the power plant of choice for aircraft applications.

BritainAir Commodore Sir Frank Whittle (1 June 1907 9 August 1996) was an engineer of the English Royal Air Force credited due to the first design of a turbo. A patent was submitted by Maxime Guillaume in 1921 (French) for a similar invention; however, this was technically unfeasible at the time

Three years of his work were spent in designing an appropriate combustion chamberEarly Whittler vaporizer combustorThe engine employed 10 separate tubular combustors in a reverse-flow arrangement to permit a short engine shaftMore than 30 configurations were evaluatedfuel was heated in tubes located in the flame zone at high pressure to prevent vaporisation of fuel before its injectionThermal cracking, coking up of the vaporizer tubes and the difficulties in controlling the fuel flow rate were the main drawbacks

Early Whittle atomizer combustor.A pressure-swirl atomizer is implemented with a wide spray cone anglemost of the primary-zone airflow enters the combustion zone through a large air swirler located at the upstream end of the liner around the fuel nozzleThis arrangement not only anchored the flame, but also provided the rapid mixing of fuel vapor, air, and combustion productsStub pipes are used to inject air to complete the combustion and reduce the air temperature

Whittle atomizer combustor implemented in the jet w1 The Jet W1 did the first British turbojet-powered flight on the evening of May 15, 1941

the De Havilland GoblinIt is the first British engine to use straight-through combustors, designed by Frank HalfordIt is implemented to power the de Havilland Vampire on the 26th of September of 1943It is implemented to power the Lockheed P-40 in its early designs. Later on the engine is replaced by the General Electrics I-40 engine that provides 33% more thrust

the De Havilland Goblin

The de Havilland DH.100 Vampire (1945)

(End of 1940)MetropolitanVickers Beryl engineThis engine implements the first British annular combustorA noteworthy feature of this combustor is the use of upstream fuel injectionThe immersion of key components in the flame is it main drawbackCooling arrangements for the atomizer feed arm could be provided, but it was difficult to eliminate entirely the problem of carbon deposition on the atomizer faceDilution air is injected through rows of scoops

MetropolitanVickers Beryl engineThe first row of scoops provides air for the completion of combustion, with any excess serving as dilution airThe second row of scoops is solely for dilution purposesLow pressure loss and low pattern factorThe scoops are prone to burnout, since they are in contact with burned gases at elevated velocitiesit carries a high weight penalty

GermanyHans Joachim Pabst von Ohain (14 December 1911 13 March 1998) was a German engineer, and designer of the first operational jet engine.[1] His first design ran in March 1937, and it was one of his engines that powered the first all-jet aircraft, the prototype of the Heinkel He 178 in late August 1939. In spite of these early successes, other German designs quickly eclipsed von Ohain's, and none of his engine designs entered widespread production or operational use.

Jumo 004 Germany - Junkers Flugzeug und Motorenwerke

it was the worlds first mass-produced turbojet and one that saw extensive service in World War IIIt was among the first engines to employ axial flow turbo-machinery and straight-through combustorsEach of the six tubular combustors was supplied with fuel at pressures up to 5.2 MPa (750 psi) from a pressure-swirl atomizer, which sprayed the fuel upstream into the primary combustion zoneJumo 004 Germany - Junkers Flugzeug und MotorenwerkeIt was the worlds first mass-produced turbojet and one that gave extensive service in the World War IIIt was among the first engines to employ axial flow turbo-machinery and straight-through combustorsEach of the six tubular combustors was supplied with fuel at pressures up to 5.2 MPa (750 psi) from a pressure-swirl atomizer, which sprayed the fuel upstream into the primary combustion zone

Jumo 004 Germany - Junkers Flugzeug und MotorenwerkeThe primary air flowed into the liner through six swirl vanes, the amount of air being sufficient to achieve near-stoichiometric combustion at the engine design pointMixing between combustion products and dilution air was achieved using an assembly of stub pipes that were welded to a ring at their upstream end and to the outer perimeterCombustion products flowed radially outward through the gaps between the stub pipes to meet and mix with part of the cold secondary airThe remaining secondary air flowed through the stub pipes, incidentally serving to protect them from burnout because of their immersion in the hot combustion gases, to provide further mixing of hot and cold gases in the recirculation zone created by the presence of the baffle.

BMW 003It is also produced during the World War II Annular combustor fitted with 16 equally spaced, downstream spraying, pressure atomizersEach fuel nozzle is surrounded by a baffle and the primary combustion air flows both through and around itDilution air flowed through 40 scoops attached to the outer liner, alternating in circumferential locations with 40 similar scoops attached to the inner liner.The end result is a combustor having a relatively low pressure loss, but also a fairly high length/height ratio

United STATEsUSA could not produce a design of its own jet engines till much after the Second World War. During later years of war, Britain supplied details of their engines that started jet engine era in USAGE was tasked by the US Government to produce Jet Engines based on British (Whittles) designs and developed J31 and the J33 that later powered Lockheeds F-80 fighter and T33 (first jet trainer)GEs j47 was first US designed jet engine powering F86 and several other aircraft in late 40s and early 50sPratt and Whitney (P&W) purchased the RR Nene license to produce their first jet engine j42Another engine, produced by Wright as j-65 was the Armstrong Siddelys Saphhire to power the B-57 light bomber also a licensed production of British CanberraW2 engineThe W2B engine is designated to produce 1600 lbf of thrust, while the W1 was able to produce only 850lbf of thrustIn 1941, a W2B engine, complete with drawings, is delivered to the General Electric Company (GE)GE produces two engines during the next six months when Pratt and Whitney (P&W) license the Nene engine from Rolls Royce , able to produce GE and P&W lost no time in producing their own independent combustor designs (J 31, J33, J35, J47)

J-33 (1943 in USA) & RR Nene J33Maximum thrust: 4,600 lbf (20.46 kn) at 11,500 rpm at sea level for take-offNormal thrust, static: 3,900 lbf (17.35 kn) at 11,000 rpm at sea levelOverall pressure ratio: 4.1:1turbine inlet temperature: 1,320 f (989 k; 716 c)NENEMaximum thrust: 5,000 lbf (22.24 kN) at 12,300 rpm at sea level for take offMax. cruising, static: 4,360 lbf (19.39 kN) at 12,000 rpm at sea levelIdling, static: 120 lbf (0.53 kN) at 2,500 rpm at sea levelCompressor pressure ratio: 4:1.Thrust-to-weight ratio: 3.226 lbf/lb (0.0315 kN/kg)

J30 (1943) and J34 (late 1940s)In 1943 Westinghouse develops successfully an axial-flow turbojet engines without any European inputAn annular combustor is selected for the J30 engineA dual-annular configuration is adopted for the J34The dual-annular concept is ahead of its time so it is discarded in future gas turbines until 1970, when GE implements this design to reduce emissions for their CFM56-B engine

J57 (1953)Maximum thrust: 12030 lbf (53.5 kN) @ Take-off, SLS, ISAOverall pressure ratio: 12.5:1Air mass flow: 180 lb/s (81.65 kg/s)Thrust-to-weight ratio: 3.44

J57 (1953)P&W employs eight tubular liners located within an annular casingEach liner has a perforated tube along its central axis that extended about halfway down the linerThe central tube converts the tubular liner into a small annular combustorSix equally spaced pressure-swirl nozzles supply the fuel

By the end of the 1940s, the development work carried out in the UK, Germany, and the United States had established the basic design features of aero-engine combustors that have remained largely unchangedBasic design featuresThis simple arrangement is impractical because the pressure loss incurred would be excessive (Vel = 170 m/s Dp could be 1/3 the gain of pressure)A flow reversal must be created to provide a low-velocity region in which to anchor the flameA liner to produce the desired temperature rise is necessaryRepresentation of the desired combustion chamber

Basic design featuresThe overall chamber air/fuel ratio must normally be around 3040, which is well outside the limits of flammability for hydrocarbonair mixturesIdeally, the air/fuel ratio in the primary combustion zone should be around 18, although higher values (around 24) are sometimes preferred if low emissions of nitric oxides (NOX) is a prime considerationThe choice of a particular type and layout of combustion chamber is determined largely by engine specifications, but it is also strongly influenced by the desirability of using the available space as effectively as possibleOn large aircraft engines, the chamber is almost invariably of the straight-through typeIn smaller engines, the reverse-flow annular combustor provides a more compact unit and allows close coupling between the compressor and turbineCombustor RequirementsHigh-combustion efficiencyReliable and smooth ignitionWide stability limitsLow pressure lossAn outlet temperature distribution able to maximize the lives of the turbine blades and nozzle guide vanesLow emissions of smoke and gaseous pollutant speciesFreedom from pressure pulsations and other manifestations of combustion-induced instabilitySize and shape compatible with engine envelopeDesign for minimum cost and ease of manufacturingMaintainabilityDurabilityPetroleum, synthetic, and biomass-based multi-fuel capability (Power Generation)size and weight (aircraft)

COMBUSTORS TYPE

Tubular (caN-type) combustorA cylindrical liner mounted concentrically inside a cylindrical casingMost of the early jet engines feature this type of combustor usually in numbers varying from 6 to 16 per engineRelatively little time and money is incurred in their development Their excessive length and weight prohibit their use in new aircraft enginesTheir main application is to industrial units where accessibility and ease of maintenance are prime considerations

Tuboannular combustorIn general, a group of tubular liners, usually from 6 to 10, is arranged inside a single annular casingThis concept attempts to combine the compactness of the annular chamber with the mechanical strength of the tubular chamberThis concept needs for interconnectors (cross-fire tubes) such as tubular chambersPrototypes are easy to evaluate, with very modest air supply, when compared with annular combustion chambers (using only a small segment of the whole chamber)Its drawbacks emerge when trying to achieve a satisfactory and consistent airflow pattern; in particular, the design of the diffuser can present serious difficulties.

annular combustorAn annular liner is mounted concentrically inside an annular casingit is an ideal form of chamber, because its clean aerodynamic layout results in a compact unit of lower pressure loss than other combustor typesThe heavy buckling load on the outer liner is it main drawback, so it is initially implemented for low-pressure-ratio gas turbines Testing of new prototypes is another drawbackBy the 1960s, the annular layout was firmly established as the automatic choice for all new aircraft engines.

Diffuser The cold loss represents the sum of the losses arising in the diffuser and the linerPressure loss in the diffuser is entirely wastedthe pressure drop across the liner wall is manifested as turbulence, which is highly beneficial to both combustion and mixingIt is customary to use a diffuser to recover the kinetic energy

Diffuser A relatively long aerodynamic diffuser to achieve maximum recovery of dynamic pressure is implemented in case (a)It typically recovers about 35%before the air reaches the liner it divides and flows into three separate diffusing passagesThe dump or step diffuse is shown in figure (b)Short conventional diffuser in which the air velocity is reduced to almost half its inlet valueDump diffusers are now generally preferred they operate better at different conditions and they reduce the weight and size of the engine

PRIMARY ZONE It creates toroidal flow patterns through swirl vanes and/or holes drilled in the liner wall to inject airIt provides wide stability limits, good ignition performance, and freedom from the type of flow instabilities that often give rise to combustion pulsations and noise

Intermediate ZoneDropping the temperature to an intermediate level by the addition of small amounts of air encourages the burnout of soot and allows the combustion of CO and any other unburned hydrocarbons (UHC) to proceed to completionBy around 1970, the increment of the compressor pressure ratio forced the traditional form of intermediate zone to disappeared since more air was required for combustion and liner-wall cooling However, the desirability of an intermediate zone remains

Dilution ZoneIt provides an outlet stream with a temperature distribution that is acceptable to the turbine pattern factorThe amount of air available for dilution is usually between 20 and 40% of the total combustor airflowThe length/diameter ratios of dilution zones all tend to lie in a narrow range between 1.5 and 1.8