Rolls-Royce Gem 60 | Rolls-Royce AJ65 Avon

Rolls-Royce Gem 60

Rolls-Royce AJ65 Avon

Work commenced early in 1945 on the AJ65 axial flow turbo-jet with a design thrust of 6500 lb. This figure was reached in 1951 with the 100 series RA3. In 1953 the considerably redesigned 200 series RA14 was type tested at 9500 lb thrust. Development culminated in the 300 series RB146 which produced 17.110lb thrust with afterburning

Pawan Hans Helicopter Limited Aviation Officer Jobs 2012

Pawan Hans Helicopter Limited (PHHL) a Government of India Enterprise (An ISO 9001:2008, ISO 14001:2004 & OHAS 18001:2007 Company) invites application for the post of Aviation Officer on Contract Basis.

Post Name	Upper Age Limit	Qualification & Experience
AVIATION OFFICER	30 Years	Retired Air Force / Army / Naval Officer in possession of CPL/CHPL with Aviation background and adequate experience of Flying Operations and Flight Safety.  The incumbent will be assisting the Operations Department in liaisoning with various Civil / Govt. agencies in the matter of flying operations and responsible for seaplane operations and all related returns and records as per Government/DGCA Guidelines.

Walk In Interview Date: 25th January 2012

Registration Time: 10.30 AM to 12.30 PM

Venue: PAWAN HANS HELICOPTERS LIMITED, C-14, SECTOR-1, NOIDA - 201 301

Desirous candidates may download application blank from the website www.pawanhans.co.in and bring it duly filled in affixing a recent passport size photograph accompanied with copies of testimonials in support of age, caste/class, qualification & experience etc and Demand Draft of Rs. 250/- (Rupees two hundred and fifty only) drawn in favour of Pawan Hans Helicopters Ltd  and payable at Delhi/Noida (SC & ST candidates are exempted from payment of application fee).  The candidates must also bring their original testimonials for verification to the interview.

Detailed Info: https://www.pawanhans.co.in/Site/CareerDetails.aspx?id=74

Middle East short of skilled aviation personnel - NEWS

The impact of the Middle East aviation sector’s achievements to date and its plans for the future have sent shockwaves to all corners of the global industry.

One of the major challenges of this dramatic growth has been the ability to attract and, more importantly, retain the right numbers of appropriately skilled people. The war for talent has affected all employees, but especially highly-skilled personnel like pilots and engineers.

GOVERNMENT OF INDIA DEPARTMENT OF SPACE | ADV NO.SAC:01/2012

 

GOVERNMENT OF INDIA

DEPARTMENT OF SPACE

INDIAN SPACE RESEARCH ORGANISATION 

SPACE APPLICATIONS CENTRE 

Ahmedadad - 380015

www.sac.gov.in

Fuel system - INTRODUCTION

INTRODUCTION
1. The functions of the fuel system are to provide the engine with fuel in a form suitable for combustion  and to control the flow to the required quantity necessary for easy starting, acceleration and stable running, at all engine operating conditions. To do this, one or more fuel pumps are used to deliver the fuel to the fuel spray nozzles, which inject it into the combustion system (Part 4) in the form of an atomized spray. Because the flow rate must vary according to the amount of air passing through the engine to maintain a constant selected engine speed or pressure ratio, the controlling devices are fully automatic with the exception of engine power selection, which is achieved by a manual throttle orpower lever. A fuel shut-off valve (cock) control lever is also used to stop the engine, although in some instances these two manual controls are combined for single-lever operation.
2. It is also necessary to have automatic safety controls that prevent the engine gas temperature, compressor delivery pressure, and the rotating assembly speed, from exceeding their maximum limitations.
3. With the turbo-propeller engine, changes in propeller speed and pitch have to be taken into account due to their effect on the power output of the engine. Thus, it is usual to interconnect the throttle lever and propeller controller unit, for by so doing the correct relationship between fuel flow and airflow is maintained at all engine speeds and the pilot is given single-lever control of the engine. Although the maximum speed of the engine is normally determined by the propeller speed controller, over-speeding is ultimately prevented by a governor in the
fuel system.
4. The fuel system often provides for ancillary functions, such as oil cooling (Part 8) and the hydraulic control of various engine control systems; for example, compressor airflow control (Part 3).

Pressure control (turbo-jet engine)

Pressure control (turbo-jet engine)
17. In the pressure control system illustrated in fig. 10-5, the rate of engine acceleration is controlled by
a dashpot throttle unit. The unit forms part of the fuel control unit and consists of a servo-operated throttle,
which moves in a ported sleeve, and a control valve.

        The control valve slides freely within the bore of the throttle valve and is linked to the pilot's throttle by a rack and pinion mechanism. Movement of the throttle lever causes the throttle valve to progressively uncover ports in the sleeve and thus increase the fuel flow. Fig. 10-6 shows the throttle valve and control valve in their various controlling positions.

                                                     

18. At steady running conditions, the dashpot throttle valve is held in equilibrium by throttle servo pressure opposed by throttle control pressure plus spring force. The pressures across the pressure drop control diaphragm are in balance and the pump servo pressure adjusts the fuel pump to give a constant fuel flow.

19. When the throttle is opened, the control valve closes the low pressure (L.P.) fuel port in the sleeve
and the throttle servo pressure increases. The throttle valve moves towards the selected throttle position until the L.P. port opens and the pressure balance across the throttle valve is restored. The decreasing fuel pressure difference across the throttle valve is sensed by the pressure drop control diaphragm, which closes the spill valve to increase the pump servo pressure and therefore the pump output. The spill valve moves into the sensitive position, controlling the pump servo mechanism so that the correct fuel flow is maintained for the selected throttle position.

Fuel control system

Fuel system





hydro-mechanical, and the acceleration and speed control and pressure ratio control systems, which are mechanical. With the exception of the pressure ratio control system, which uses a gear-type pump, all the systems use a variable-stroke, multi-plunger type fuel pump to supply the fuel to the spray nozzles.

8. Some engines are fitted with an electronic system of control and this generally involves the use of electronic circuits to measure and translat changing engine conditions to automatically adjust the fuel pump output. On helicopters powered by gas turbine engines using the free-power turbine principle (Part 5), additional manual and automatic controls on the engine govern the free-power turbine and, consequently, aircraft rotor speed.

Ring seals

Ring seals

17. A ring seal (fig. 9-7) comprises a metal ring which is housed in a close fitting groove in the static housing. The normal running clearance between the ring and rotating shaft is smaller than that which can be obtained with the labyrinth seal. This is because the ring is allowed to move in its housing whenever the shaft comes into contact with it.
18. Ring seals are used for bearing chamber sealing, except in the hot areas where oil degradation due to heat would lead to ring seizure within its housing.

Hydraulic seals

19. This method of sealing is often used between two rotating members to sea a bearing chamber.
Unlike the labyrinth or ring seal, it does not allow a controlled flow of air to traverse across the seal,

20. Hydraulic seals (fig. 9-7) are formed by a seal fin immersed in an annulus of oil which has been created by centrifugal forces. Any difference in air pressure inside and outside of the bearing chamber is compensated by a difference in oil level either side of the fin.

Carbon seals

21. Carbon seals (fig. 9-7) consist of a static ring of carbon which constantly rubs against a collar on a rotating shaft. Several springs are used to maintain contact between the carbon and the collar. This type of seal relies upon a high degree of contact and does not allow oil or air leakage across it. The heat caused by friction is dissipated by the oil system.

Brush seals

SEALING, A hypothetical turbine cooling and sealing arrangement,A generator cooling system,Labyrinth seals.

Fig. 9-5 A hypothetical turbine cooling and sealing arrangement.

SEALING :
12. Seals are used to prevent oil leakage from the engine bearing chambers, to control cooling airflows

and to prevent ingress of the mainstream gas into the turbine disc cavities.
13. Various sealing methods are used on gas turbine engines. The choice of which method is
dependent upon the surrounding temperature and pressure, wearability, heat generation, weight, space
available, ease of manufacture and ease of installa- tion and removal. Some of the sealing methods are
described in the following paragraphs. A hypothetical turbine showing the usage of these seals is shown in
fig. 9-5.

Bearing chamber cooling, Development of high pressure turbine blade cooling,High pressure nozzle guide vane construction and cooling ,Accessory cooling

Fig. 9-3 Development of high pressure turbine blade cooling.

Bearing chamber cooling


9. Air cooling of the engine bearing chambers is not normally necessary since the lubrication system
(Part 8) is adequate for cooling purposes.

Turbine cooling,Nozzle guide vane and turbine blade cooling arrangement

Turbine cooling :


5. High thermal efficiency is dependent upon high turbine entry temperature, which is limited by the
turbine blade and nozzle guide vane materials. Continuous cooling of these components allows their

Fig. 9-2 Nozzle guide vane and turbine blade cooling arrangement.

environmental operating temperature to exceed the material's melting point without affecting the blade
and vane integrity. Heat conduction from the turbine blades to the turbine disc requires the discs to be
cooled and thus prevent thermal fatigue and uncon- trolled expansion and contraction rates.

A centrifugal breather, A thread-type oil filter,A typical pressure and scavenge filter,LUBRICATING OILS

Fig. 8-11 A thread-type oil filter.

27. In some engines, to minimize the effect of the dynamic loads transmitted from the rotating
assemblies to the bearing housings, a 'squeeze film' type of bearing is used (fig. 8-9). They have a small
clearance between the outer race of the bearing and housing with the clearance being filled with oil. The
oil film dampens the radial motion of the rotating assembly and the dynamic loads transmitted to the
bearing housing thus reducing the vibration level of the engine and the possibility of damage by fatigue.

28. To prevent excessive air pressure within the oil tank, gearboxes and bearing chambers, a vent to
atmosphere is incorporated within the lubrication system. Any oil droplets in the air are separated out
by a centrifugal breather prior to the air being vented overboard. Some breathers may incorporate a
porous media, forming de-aerator segments, which improves the efficiency of the oil separation (fig, 8-10)

A low pressure fuel-cooled oil cooler,A magnetic chip detector.

Fig. 8-7  A low pressure fuel-cooled oil cooler.

Fig. 8-8 A magnetic chip detector.

25. The air-cooled oil cooler is similar to the fuel-cooled type in both construction and operation; the
main difference is that air is used as the cooling medium.

Principle of a gear pump

20. The most common type of oil distribution device is a simple orifice which directs a metered amount of
oil onto its target. These jet orifices are positioned as close to the target area as possible to overcome the
possibility of the local turbulent environment deflecting the jet of oil. The smallest diameter of a jet
orifice is 0.04 inch which allows a flow of 12 gallons per hour when operating at a pressure of 40 lb. per
sq. in. The use of restrictors upstream can reduce the flow rate if required.

Fig. 8-5 Principle of a gear pump.

21. All engines transfer heat to the oil by friction, churning and windage within a bearing chamber or
gearbox. It is therefore common practice to fit an oil cooler in recirculatory oil systems. The cooling
medium may be fuel or air and, in some instances, both fuel-cooled and air-cooled coolers are used.

22. Some engines which utilize both types of cooler may incorporate an electronic monitoring system

which switches in the air-cooled cooler only when it is necessary. This maintains the ideal oil temperature

and improves the overall thermal efficiency. 

An oil tank,OIL SYSTEM COMPONENTS

OIL SYSTEM COMPONENTS :


12. The oil tank (fig. 8-4) is usually mounted on the engine and is normally a separate unit although it
may also be an integral part of the external gearbox. It must have provision to allow the lubrication system
to be drained and replenished. A sight glass or a system contents to be checked. The filler can be

Fig. 8-4 An oil tank.

either the gravity or pressure filling type; on some engines both types are fitted. Provision is also made
for a continuous supply of oil to be made available in aircraft which are designed to operate during
inverted flight conditions. Since air is mixed with the oil in the bearing chambers, a de-aerating device is
incorporated within the oil tank which removes the air from the returning oil.

A total loss (expendable) oil system,Total loss (expendable) system :

Total loss (expendable) system :

Fig. 8-3 A total loss (expendable) oil system.

10. For engines which run for periods of short duration, such as booster and vertical lift engines,the total loss oil system is generally used. The system is simple and incurs low weight penalties
because it requires no oil cooler, scavenge pump or filters. On some engines oil is delivered in a
continuous flow to the bearings by a plunger-type pump, indirectly driven from the compressor shaft; on
others it is delivered by a piston-type pump operated by fuel pressure (fig. 8-3).

Full flow system,A full flow type oil system.

Full flow system

Fig. 8-2 A full flow type oil system.

7. Although the pressure relief valve system operates satisfactorily for engines which have a low
bearing chamber pressure, which does not unduly increase with engine speed, it becomes an
undesirable system for engines which have high chamber pressures. For example, if a bearing
chamber has a maximum pressure of 90 lb. per sq. in. It would require a pressure relief valve setting of
130 lb. per sq. in. to produce a pressure drop of 40 lb. per sq. in. at the oil feed jet. This results in theneed for large pumps and difficulty in matching the required oil flow at slower speeds.

Accessory drives

Accessory drives:



INTRODUCTION
1. Accessory units provide the power for aircraft hydraulic, pneumatic and electrical systems in  addition to providing various pumps and control systems for efficient engine operation. The high level  of dependence upon these units requires an extremely reliable drive system.taken from a rotating engine shaft, via an internal  mount for the accessories and distributes the starter may also be fitted to provide an input torque   by-pass engine takes between 400 and 500
2. The drive for the accessory units is typically gearbox, to an external gearbox which provides a appropriate geared drive to each accessory unit. A to the engine. An accessory drive system on a high  horsepower from the engine.

Fig. 7-1 Mechanical arrangement of accessory drives.

Internal air system,General internal airflow pattern,COOLING

 Internal air system :



INTRODUCTION

Fig. 9-1 General internal airflow pattern.

1. The engine internal air system is defined as those airflows which do not directly contribute to the engine thrust. The system has several important functions to perform for the safe and efficient operation of the engine. These functions include internal engine and accessory unit cooling, bearing chamber sealing prevention of hot gas ingestion into the turbine disc cavities, control of bearing axial loads, control of turbine blade tip clearances (Part 5) and engine anti-icing (Part 13). The system also supplies air for the aircraft services. Up to one fifth of the total engine core mass airflow may be used for these various functions.

Gear train drive ,Gear train drive ,External gearbox,Auxiliary gearbox

Gear train drive :
13. When space permits, the drive may be taken to the external gearbox via a gear train (fig. 7-1). This involves the use of spur gears, sometimes incorpo-rating a centrifugal breather (Part 8). However, it is rare to find this type of drive system in current use.
Intermediate gearbox
14. Intermediate gearboxes are employed when it is not possible to directly align the radial driveshaft with
the external gearbox. To overcome this problem an intermediate gearbox is mounted on the high pressure compressor case and re-directs the drive, through bevel gears, to the external gearbox. An example of this layout is shown in fig. 7-1.

A pressure relief valve type oil system,Pressure relief valve system,INTRODUCTION,Lubrication

Lubrication :

INTRODUCTION :

Fig. 8-1 A pressure relief valve type oil system.

1. The lubrication system is required to provide lubrication and cooling for all gears, bearings and

splines. It must also be capable of collecting foreign matter which, if left in a bearing housing or gearbox,

can cause rapid failure. Additionally, the oil must protect the lubricated components which are manu-

fractured from non-corrosion resistant materials. The oil must accomplish these tasks without significant

deterioration.

2. The requirements of a turbo-propeller engine are somewhat different to any other types of aero gas
turbine. This is due to the additional lubrication of the heavily loaded propeller reduction gears and the
need for a high pressure oil supply to operate the propeller pitch control mechanism.

3. Most gas turbine engines use a self-contained re circulatory lubrication system in which the oil is
distributed around the engine and returned to the oil tank by pumps. However, some engines use a
system known as the total loss or expendable system in which the oil is spilled overboard after the engine
has been lubricated.

LUBRICATING SYSTEMS

Gearbox sealing,Materials,An external gearbox with auxiliary gearbox drive

Gearbox sealing

Fig. 7-5 An external gearbox with auxiliary gearbox drive.

26. Sealing of the accessory drive system is primarily concerned with preventing oil loss. The
internal gearbox has labyrinth seals where the static casing mates with the rotating compressor shaft. For
some o! the accessories mounted on the external gearbox, an air blown pressurized labyrinth seal is employed. This prevents oil from the gearbox entering the accessory unit and also prevents con-tamination of the gearbox, and hence engine, in the event of an accessory failure. The use of an air blown
seal results in a gearbox pressure of about 3 lbs. per sq. in. above atmospheric pressure. To supplement a
labyrinth seal, an 'oil thrower ring' may be used. This involves the leakage oil running down the driving
shaft and being flung outwards by a flange on the rotating shaft. The oil is then collected and returned
to the gearbox.

CONSTRUCTION AND MATERIALS,Gears

CONSTRUCTION AND MATERIALS
Gears


23. The spur gears of the external or auxiliary gearbox gear train (fig. 7-4 and 7-5) are mounted
between bearings supported by the front and rear casings which are bolted together. They transmit the

Fig. 7-4 An external gearbox and accessory units.

drive to each accessory unit, which is normally between 5000 and 6000 r.p.m. for the accessory
units and approximately 20,000 r.p.m. for the centrifugal breather.

24. All gear meshes are designed with 'hunting tooth' ratios which ensure that each tooth of a gear
does not engage between the same set of opposing teeth on each revolution. This spreads any wear
evenly across all teeth.

An internal gearbox,Direct drive

Fig. 7-3 An internal gearbox.

Direct drive

11. In some early engines, a radial driveshaft was used to drive each, or in some instances a pair, of

accessory units. Although this allowed each accessory unit to be located in any desirable location

around the engine and decreased the power transmitted through individual gears, it necessitated

a large internal gearbox. Additionally, numerous radial driveshafts had to be incorporated within the

design. This led to an excessive amount of time required for disassembly and assembly of the engine

for maintenance purposes.

Mechanical arrangement of internal gearboxes.

6. To minimize unwanted movement between the compressor shaft bevel gear and radial driveshaft

(fig. 7-1)

bevel gear, caused by axial movement of the compressor shaft, the drive is taken by one of three basic methods (fig. 7-2). The least number of components is used when the compressor shaft bevel gear is mounted as close to the compressor shaft location bearing as possible, but a small amount of movement has to be accommodated within the meshing of the bevel gears. Alternatively, the compressor shaft bevel gear may be mounted on a stub shaft which has its own location bearing. The stub shaft is splined onto the compressor shaft which allows axial movement without affecting the bevel gear mesh. A more complex system utilizes an idler gear which meshes with the compressor shaft via straight spur gears, accommodating the axial movement, and drives the radial driveshaft via a bevel gear arrangement. The latter method was widely employed on early engines to overcome gear engagement difficulties at high speed.

Rolls-Royce Gnome

De Havilland H2 Ghost

                             

                                            Rolls-Royce Gnome

The Ghost was designed as a larger and more powerful version of the Goblin. After running for the first time on 2 September 1945 the engine was cleared for flight in the outer nacelles of an Avro Lancastrian at 4000 lb thrust. The Ghost later went into production at 5000 lb thrust to power the de Havilland Comet 1 airliner and Venom fighter

EXHAUST SYSTEM CONSTRUCTION AND MATERIALS

Fig. 6-6 An insulating blanket.

CONSTRUCTION AND MATERIALS
13. The exhaust system must be capable of withstanding the high gas temperatures and is therefore manufactured from nickel or titanium. It is also necessary to prevent any heat being transferred to the surrounding aircraft structure. This is achieved by passing ventilating air around the jet pipe, or by lagging the section of the exhaust system with an insulating blanket (fig. 6-6). Each blanket has an inner layer of fibrous insulating material contained by an outer skin of thin stainless steel, which is dimpled to increase its strength. In addition, acousticallyabsorbent materials are sometimes applied to the exhaust system to reduce engine noise (Part 19).

Exhaust system | EXHAUST GAS FLOW

Fig. 6-1 A basic exhaust system.

INTRODUCTION
1. Aero gas turbine engines have an exhaust system which passes the turbine discharge gases to atmosphere at a velocity, and in the required direction, to provide the resultant thrust. The velocity and pressure of the exhaust gases create the thrust in the turbo-jet engine (para. 5) but in the turbopropeller engine only a small amount of thrust is contributed by the exhaust gases, because most of the energy has been absorbed by the turbine for driving the propeller. The design of the exhaust system therefore, exerts a considerable influence on the performance of the engine. The areas of the jet pipe and propelling or outlet nozzle affect the turbine entry temperature, the mass airflow and the velocity and pressure of the exhaust jet.
2. The temperature of the gas entering the exhaust system is between 550 and 850 deg. C. according to the type of engine and with the use of afterburning (Part 16) can be 1,500 deg. C. or higher. Therefore, it is necessary to use materials and a form of construction that will resist distortion and cracking, and prevent heat conduction to the aircraft structure.
3. A basic exhaust system is shown in fig. 6-1. The use of a thrust reverser (Part 15), noise suppressor (Part 19) and a two position propelling nozzle entails a more complicated system as shown in fig. 6-2. The low by-pass engine may also include a mixer unit (fig. 6-4) to encourage a thorough mixing of the hot and cold gas streams.

Fig. 6-2 Exhaust system with thrust reverser, noise suppressor and two position propelling nozzle.

Rolls-Royce RB211-535E4 | Rolls-Royce RB50 Trent

Rolls-Royce RB211-535E4

 

 

 

 

Late in 1943 the decision was taken at Rolls- Royce to build a turbo-prop for aircraft speeds of around 400 mph. The resulting engine, known as the RB50 Trent, was basically a Derwent II with a flexible quillshaft to reduction gear and propeller. On 20 September 1945 a Gloster Meteor, fitted with two Trents, became the world’s first turboprop powered aircraft to fly.           

Rolls-Royce RB50 Trent

Nozzle guide vanes | Turbine blades | BALANCING

Fig. 5-12 Section through a dual alloy disc.

Nozzle guide vanes
25. Due to their static condition. the nozzle guide vanes do not endure the same rotational stresses as the turbine blades. Therefore, heat resistance is the property most required. Nickel alloys are used, although cooling is required to prevent melting. Ceramic coatings can enhance the heat resisting properties and, for the same set of conditions, reduce the amount of cooling air required, thus improving engine efficiency.
Turbine discs

Nozzle guide vanes | Contra-rotating turbine | COMPRESSOR-TURBINE MATCHING

Fig. 5-8 Typical nozzle guide vanes showing their shape and location.

Nozzle guide vanes
13. The nozzle guide vanes are of an aerofoil shape with the passage between adjacent vanes forming a convergent duct. The vanes are located (fig. 5-8) in the turbine casing in a manner that allows for expansion.
14. The nozzle guide vanes are usually of hollow form and may be cooled by passing compressor delivery air through them to reduce the effects of high thermal stresses and gas loads. For details of turbine cooling, reference should be made to Part 9.

Fig. 5-9 Various methods of attaching blades to turbine discs.

15. Turbine discs are usually manufactured from a machined forging with an integral shaft or with a flange onto which the shaft may be bolted. The disc also has, around its perimeter, provision for the attachment of the turbine blades.
16. To limit the effect of heat conduction from the turbine blades to the disc a flow of cooling air is passed across both sides of each disc (Part 9). Turbine blades
17. The turbine blades are of an aerofoil shape, designed to provide passages between adjacent blades that give a steady acceleration of the flow up to the ’throat’, where the area is smallest and the velocity reaches that required at exit to produce the required degree of reaction (para. 5).
18. The actual area of each blade cross-section is fixed by the permitted stress in the material used and by the size of any holes which may be required for cooling purposes (Part 9). High efficiency demands thin trailing edges to the sections, but a compromise has to be made so as to prevent the blades cracking due to the temperature changes during engine operation.
19. The method of attaching the blades to the turbine disc is of considerable importance, since the stress in the disc around the fixing or in the blade root has an important bearing on the limiting rim speed. The blades on the early Whittle engine were attached by the de Laval bulb root fixing, but this design was soon superseded by the ’fir-tree’ fixing that is now used in the majority of gas turbine engines. This type of fixing involves very accurate machining to ensure that the loading is shared by allthe serrations. The blade is free in the serrations when the turbine is stationary and is stiffened in the root by centrifugal loading when the turbine isrotating. Various methods of blade attachment are shown in fig. 5-9; however, the B.M.W. hollow blade and the de Laval bulb root types are not nowgenerally used on gas turbine engines.

ENERGY TRANSFER FROM GAS FLOW TO TURBINE

Fig. 5-6 A typical turbine blade showing
twisted contour.

ENERGY TRANSFER FROM GAS FLOW TO TURBINE
6. From the description contained in para. 1, it will be seen that the turbine depends for its operation on the transfer of energy between the combustion gases and the turbine. This transfer is never 100 per cent because of thermodynamic and mechanicallosses, (para. 11). 

Fig. 5-7 Gas flow pattern through nozzle and blade

  7.when the gas is expanded by the combustion process (Part 4), it forces its way into the discharge nozzles of the turbine where, because of theirconvergent shape, it is accelerated to about the speed of sound which, at the gas temperature, isabout 2,500 feet per second. At the same time thgas flow is given a ’spin’ or ’whirl’ in the direction o rotation of the turbine blades by the nozzle guid vanes. On impact with the blades and during th subsequent reaction through the blades, energy is absorbed, causing the turbine to rotate at high speed and so provide the power for driving the turbine shaft and compressor.

Turbines

Fig. 5-1 A triple-stage turbine with single shaft system.

Fig. 5-2 A twin turbine and shaft arrangement.

INTRODUCTION
1. The turbine has the task of providing the power to drive the compressor and accessories and, in the case of engines which do not make use solely of a jet for propulsion, of providing shaft power for a propeller or rotor. It does this by extracting energy from the hot gases released from the combustion system and expanding them to a lower pressure and temperature. High stresses are involved in this process, and for efficient operation, the turbine blade tips may rotate at speeds over 1,500 feet per second, The continuous flow of gas to which the turbine is exposed may have an entry temperature between 850 and 1,700 deg. C. and may reach a velocity of over 2,500 feet per second in parts of the turbine.
2. To produce the driving torque, the turbine may consist of several stages each employing one row of stationary nozzle guide vanes and one row of moving blades (fig. 5-1). The number of stages depends upon the relationship between the power required from the gas flow, the rotational speed at which it must be produced and the diameter of turbine permitted.
3. The number of shafts, and therefore turbines, varies with the type of engine; high compression ratio engines usually have two shafts, driving high and low pressure compressors (fig, 5-2). On high by-pass ratio fan engines that feature an intermediate pressure system, another turbine may be interposed between the high and low pressure turbines, thus forming a triple-spool system (fig, 5-3). On some engines, driving torque is derived from a free-power turbine (fig. 5-4). This method allows the turbine to run at its optimum speed because it is mechanically independent of other turbine and compressor shafts.

Rolls-Royce Turbomeca Adour Mk102 | Rolls-Royce RB37 Derwent V

Rolls-Royce RB37 Derwent V

Rolls-Royce Turbomeca Adour Mk102

Work commenced in January 1945 on a 0.855 scale Nene, reduced to fit the engine nacelle of a Gloster Meteor. Known as the Derwent V the engine passed a 100 hr test at 2600 lb thrust in June 1945 and in September went into production with a service rating of 3500 lb. Two world speed records were set by Meteor IV’s powered by special Derwent V’s in November 1945 and September 1946.

Combustion stability

Fig. 4-11 Combustion stability limits.

28. Combustion stability means smooth burning and the ability of the flame to remain alight over awide operating range.
29. For any particular type of combustion chamber there is both a rich and weak limit to the air/fuel ratio, beyond which the flame is extinguished. An extinction is most likely to occur in flight during a glide or dive with the engine idling, when there is ahigh airflow and only a small fuel flow, i.e. a veryweak mixture strength.
30. The range of air/fuel ratio between the rich and weak limits is reduced with an increase of air velocity, and if the air mass flow is increased beyond a certain value, flame extinction occurs. A typical stability loopis illustrated in fig. 4-11

COMBUSTION CHAMBER PERFORMANCE

Fig. 4-9 Annular combustion chamber.

COMBUSTION CHAMBER PERFORMANCE
23. A combustion chamber must be capable of allowing fuel to burn efficiently over a wide range of operating conditions without incurring a large pressure loss. In addition, if flame extinction occurs, then it must be possible to relight. In performing these functions, the flame tube and spray nozzleatomizer components must be mechanically reliable.
24. The gas turbine engine operates on a constant pressure cycle, therefore any loss of pressure during the process of combustion must be kept to a minimum. In providing adequate turbulence and mixing, a total pressure loss varying from about 3 to 8 per cent of the air pressure at entry to the chamber is incurred.