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.

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 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.

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.

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.

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.

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 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.

Principles of operation - Pressure and velocity changes through an axial compressor.

Fig. 3-9 Pressure and velocity changes
through an axial compressor.

19. During operation the rotor is turned at high speed by the turbine so that air is continuously induced into the compressor, which is then accelerated by the rotating blades and swept rearwards onto the adjacent row of stator vanes. The pressure rise results from the energy imparted to the air in the rotor which increases the air velocity. The air is then decelerated (diffused) in the following stator passage and the kinetic energy translated into pressure. Stator vanes also serve to correct the deflection given to the air by the rotor blades and to present the air at the correct angle to the next stage of rotor blades. The last row of stator vanes usually act as air straighteners to remove swirl from the air prior to entry into the combustion system at a reasonably uniform axial velocity. Changes in pressure and velocity that occur in the airflow through the compressor are shown diagrammatically in fig. 3-9. The changes are accompanied by a progressive increase in air temperature as thepressure increases.
20. Across each stage the ratio of total pressures of outgoing air

THE AXIAL FLOW COMPRESSOR

Fig. 3-7 Typical axial flow compressors.

THE AXIAL FLOW COMPRESSOR
13. An axial flow compressor (fig. 3-7 and fig. 3-8 consists of one or more rotor assemblies that carry blades of airfoil section. These assemblies are mounted between bearings in the casings which incorporate the stator vanes. The compressor is a multi-stage unit as the amount of pressure increase by each stage is small; a stage consists of a row of rotating blades followed by a row of stator vanes. Where several stages of compression operate in series on one shaft it becomes necessary to vary the stator vane angle to enable the compressor to operate effectively at speeds below the design condition. As the pressure ratio is increased the incorporation of variable stator vanes ensures that the airflow is directed onto the succeeding stage of rotor blades at an acceptable angle, ref. para. 30, Airflow Control.

Fig. 3-8 Typical triple spool compressor.

14. From the front to the rear of the compressor, i.e. from the low to the high pressure end, there is a

Construction - Typical impellers for centrifugal compressors

Construction
10. The construction of the compressor centres around the impeller, diffuser and air intake system. The impeller shaft rotates in ball and roller bearings and is either common to the turbine shaft or split in the centre and connected by a coupling, which is usually designed for ease of detachment. Impellers
11. The impeller consists of a .forged, disc with integral, radially disposed vanes on one or both sides (fig. 3-5) forming convergent passages in conjunction with the compressor casing. The vanes may be swept back, but for ease of manufacture straight radial vanes are usually employed. To ease the air from axial flow in the entry duct on to the rotating impeller, the vanes in the centre of the impeller are curved in the direction of rotation. The curved sections may be integral with the radial vanes or formed separately for easier and more accurate manufacture.Diffusers
12. The diffuser assembly may be an integral part of the compressor casing or a separately attached assembly. In each instance it consists of a number of vanes formed tangential to the impeller. The vane passages are divergent to convert the kinetic energy into pressure energy and the inner edges of the vanes are in line with the direction of the resultant airflow from the impeller (fig. 3-6). The clearance between the impeller and the diffuser is an important factor, as too small a clearance will set up aerodynamic buffeting impulses that could be transferred to the impeller and create an unsteady airflow and vibration.

Fig. 3-6 Airflow at entry to diffuser.

 

 

Fig. 3-5 Typical impellers for centrifugal
compressors.

THE CENTRIFUGAL FLOW COMPRESSOR

THE CENTRIFUGAL FLOW COMPRESSOR

Fig. 3-3 Pressure and velocity changes
through a centrifugal compressor.

5. Centrifugal flow compressors have a single or double-sided impeller and occasionally a two-stage, single sided impeller is used, as on the Rolls-Royce Dart. The impeller is supported in a casing that also contains a ring of diffuser vanes. If a double-entry impeller is used, the airflow to the _rear side is reversed in direction and a plenum chamber is required. Principles of operation
6. The impeller is rotated at high speed by the turbine and air is continuously induced into the centre of the impeller. Centrifugal action causes it to flow radially outwards along the vanes to the impeller tip, thus accelerating the air and also causing a rise in pressure to occur. The engine intake duct may contain vanes that provide an initial swirl to the air entering the compressor.

Fig. 3-4 Impeller working clearance and
air leakage.

Airflow systems.

Fig. 2-5-1 Airflow systems

THE RELATIONS BETWEEN PRESSURE, VOLUME AND TEMPERATURE

THE RELATIONS BETWEEN PRESSURE, VOLUME AND TEMPERATURE 

Fig. 2-3 An airflow through divergent and convergent ducts.

7. During the working cycle of the turbine engine, the airflow or ’working fluid’ receives and gives up heat, so producing changes in its pressure, volume and temperature. These changes as they occur are closely related, for they follow a common principle that is embodied in a combination of the laws of Boyle and Charles. Briefly, this means that the product of the pressure and the volume of the air at the various stages in the working cycle is proportional to the absolute temperature of the air at those stages. This relationship applies for whatever means are used to change the state of the air. For example, whether energy is added by combustion or by compression, or is extracted by the turbine, the heat change is directly proportional to the work added or taken from the gas.