A vaporizer combustion chamber,

Fig. 4-5 A vaporizer combustion chamber

FUEL SUPPLY
12. Fuel is supplied to the airstream by one of two distinct methods. The most common is the injection of a fine atomized spray into the recirculating airstream through spray nozzles (Part 10). Thesecond method is based on the pre-vaporization ofthe fuel before it enters the combustion zone.
13. In the vaporizing method (fig.4-5) the fuel is sprayed from feed tubes into vaporizing tubes which are positioned inside the flame tube. These tubes turn the fuel through 180 degrees and, as they are heated by combustion, the fuel vaporizes before passing into the flame tube. The primary airflow passes down the vaporizing tubes with the fuel and also through holes in the flame tube entry section which provide ’fans’ of air to sweep the flame rearwards. Cooling and dilution air is metered into the flame tube in a manner similar to the atomizer flame tube.
TYPES OF COMBUSTION CHAMBER 
14. There are three main types of combustion chamber in use for gas turbine engines. These are the multiple chamber, the tubo-annular chamber and the annular chamber.
Multiple combustion chamber 
15. This type of combustion chamber is used on centrifugal compressor engines and the earlier types of axial flow compressor engines. It is a direct development of the early type of Whittle combustion chamber. The major difference is that the Whittle chamber had a reverse flow as illustrated in fig. 4-6 but, as this created a considerable pressure loss, the straight-through multiple chamber was developed by Joseph Lucas Limited.

Fig. 4-6 An early Whittle combustion chamber

Apportioning the airflow - COMBUSTION PROCESS

COMBUSTION PROCESS
4. Air from the engine compressor enters the combustion chamber at a velocity up to 500 feet per second, but because at this velocity the air speed is far too high for combustion, the first thing that the chamber must do is to diffuse it, i.e. decelerate it and raise its static pressure. Since the speed of burning kerosine at normal mixture ratios is only a few feet  per second, any fuel lit even in the diffused airstream, which now has a velocity of about 80 feet persecond, would be blown away. A region of low axialvelocity has therefore to be created in the chamber,so that the flame will remain alight throughout therange of engine operating conditions.

Fig. 4-2 Apportioning the airflow.

5. In normal operation, the overall air/fuel ratio of a combustion chamber can vary between 45:1 and 130:1, However, kerosine will only burn efficiently at, or close to, a ratio of 15:1, so the fuel must be burned with only part of the air entering the chamber, in what is called a primary combustion zone. This is achieved by means of a flame tube (combustion liner) that ha various devices for metering the airflow distribution along the chamber.

INTRODUCTION - An early combustion chamber.

INTRODUCTION

Fig. 4-1 An early combustion chamber.

1. The combustion chamber (fig. 4-1) has the difficult task of burning large quantities of fuel, supplied through the fuel spray nozzles (Part 10), with extensive volumes of air, supplied by the compressor (Part 3), and releasing the heat in such a manner that the air is expanded and accelerated to give a smooth stream of uniformly heated gas at all conditions required by the turbine (Part 5). This task must be accomplished with the minimum loss in pressure and with the maximum heat release for the limited space available.

Rolls-Royce RB41 Nene | Rolls-Royce RB211 Trent

Rolls-Royce RB41 Nene

Rolls-Royce RB211 Trent

On 17 March 1944 Rolls-Royce commenced work on the RB40 as the result of a Government request for a turbo-jet of 4200 lb thrust. After discussions with Supermarine, the airframe designers, the engine was scaled down to produce 3400 lb. The resulting Nene was eventually rated at 5000 lb and powered the Hawker Sea Hawk and Supermarine Attacker.

An electronically operated bleed valve system

Fig. 3-18 An electronically operated bleed valve system.

Fig. 3-19 Typical types of fan blades.

32. For casing designs the need is for a light but rigid construction enabling blade tip clearances to be accurately maintained ensuring the highest possible efficiency. These needs are achieved by using aluminium at the front of the compression system followed by .alloy steel as compression temperature increases. Whilst for the final stages of the compression system, where temperature requirements possibly exceed the capability of the best steel, nickel based alloys may be required. The use of titanium in .preference to aluminium and steel is now more common; particularly in military engines where its high rigidity to density ratio can result in significant weight reduction. With the development of new manufacturing methods component costs can now be maintained at a more acceptable level in spite of high initial material costs.

AIRFLOW CONTROL - A hydraulically operated bleed valve and inlet guide vane airflow control system.

Fig. 3-15 Typical variable stator vanes.

AIRFLOW CONTROL
30. Where high pressure ratios on a single shaft are required it becomes necessary to introduce airflow control into the compressor design. This may take the form of variable inlet guide vanes for the first stage plus a number of stages incorporating variable stator vanes for the succeeding stages as the shaft pressure ratio is increased (fig. 3-15). As the compressor speed is reduced from its design value these static vanes are progressively closed in order to maintain an acceptable air angle value onto the following rotor blades. Additionally interstage bleed may be provided but its use in design is now usually limited to the provision of extra margin while the engine is being accelerated, because use at steady operating conditions is inefficient and wasteful of fuel. Three types of air bleed systems are illustrated as follows: fig. 3-16 hydraulic, fig. 3-17 pneumatic and fig. 3-18 electronic.
MATERIALS
31. Materials are chosen to achieve the most cost effective design for the components in question, in practice for aero engine design this need is usually best satisfied by the lightest design that technology allows for the given loads and temperatures prevailing.

Fig. 3-16 A hydraulically operated bleed valve and inlet guide vane airflow control system.

Fig. 3-17 A pneumatically operated bleed valve system.

A typical rotor blade showing twisted contour

Fig. 3-12 A typical rotor blade showing
twisted contour.

26. The rotor blades are of airfoil section (fig. 3-12) and usually designed to give a pressure gradient along their length to ensure that the air maintains a reasonably uniform axial velocity. The higher pressure towards the tip balances out the centrifugal action of the rotor on the airstream. To obtain these conditions, it is necessary to ’twist’ the blade from root to tip to give the correct angle of incidence at each point. Air flowing through a compressor creates two boundary layers of slow to stagnant air on the inner and outer walls. In order to compensate for the slow air in the boundary layer a localized increase in blade camber both at the blade tip and root has been introduced. The blade extremities appear as ifformed by bending over each corner, hence the term’end-bend’.Stator vanes
27. The stator vanes are again of airfoil section and are secured into the compressor casing or into stator vane retaining rings, which are themselves secured to the casing (fig. 3-13). The vanes are often assembled in segments in the front stages and may be shrouded at their inner ends to minimize the vibrational effect of flow variations on the longer vanes. It is also necessary to lock the stator vanes in such a manner that they will not rotate around the casing.
OPERATING CONDITIONS
28. Each stage of a multi-stage compressor possesses

Methods of securing blades to disc and Rotors of drum and disc construction

Fig. 3-11 Methods of securing blades to disc.

Fig. 3-10 Rotors of drum and disc
construction.

23. The construction of the compressor centres around the rotor assembly and casings. The rotor shaft is supported in ball and roller bearings and coupled to the turbine shaft in a manner that allows for any slight variation of alignment. The cylindrical casing assembly may consist of a number of cylindrical casings with a bolted axial joint between each stage or the casing may be in two halves with a bolted centre line joint. One or other of these construction methods is required in order that the casing can be assembled around the rotor. Rotors

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.

Compressors - INTRODUCTION

INTRODUCTION
1. In the gas turbine engine, compression of the air before expansion through the turbine is effected by one of two basic types of compressor, one giving centrifugal flow and the other axial flow. Both types are driven by the engine turbine and are usually coupled direct to the turbine shaft.

Fig. 3-1 A typical centrifugal flow compressor.

2. The centrifugal flow compressor (fig. 3-1) is a single or two stage unit employing an impeller to accelerate the air and a diffuser to produce the required pressure rise. The axial flow compressor (fig. 3-7 and fig. 3-8) is a multi-stage unit employing alternate .rows of rotating (rotor) blades and stationary (stator) vanes, to accelerate and diffuse the air until the required pressure rise is obtained. In some cases, particularly on small engines, an axial compressor is used to boost the inlet pressure to the centrifugal.

Fig. 3-2 Specific fuel consumption and
pressure ratio.

3. With regard to the advantages and disadvantages of the two types, the centrifugal compressor is usually more robust than the axial compressor and is also easier to develop and manufacture. The axial compressor however consumes far more air than a centrifugal

Rolls-Royce RB211-22B

Rolls-Royce RB211-22B

De Havilland H1 Goblin

 Development of the de Havilland Goblin began in 1941 with the Halford H1 with a design thrust of 3000 lb. The engine passed a 25 hr special category test in September 1942 and was cleared for flight at 2000 lb thrust. This took place in a Gloster Meteor on 5 March 1943 and was also the first flight of that aircraft type.In September 1943 the first flight of a de Havilland DH100 Vampire was made with a Goblin of 2300 lb thrust.

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.

Working cycle and airflow

INTRODUCTION
1. The gas turbine engine is essentially a heat engine using air as a working fluid to provide thrust. To achieve this, the air passing through the engine has to be accelerated; this means that the velocity or kinetic energy of the air is increased. To obtain this increase, the pressure energy is first of all increased, followed by the addition of heat energy, before final conversion back to kinetic Energy in the form of a high velocity jet efflux.
WORKING CYCLE

Fig. 2-1 A comparison between the working cycle of a turbo-jet engine and a piston engine.

2. The working cycle of the gas turbine engine is similar to that of the four-stroke piston engine. However, in the gas turbine engine, combustion occurs at a constant pressure, whereas in the piston engine it occurs at a constant volume. Both engine cycles (fig. 2-1) show that in each instance there is induction, compression, combustion and exhaust.
These processes are intermittent in the case of the piston engine whilst they occur continuously in the gas turbine. In the piston engine only one stroke is utilized in the production of power, the others being involved in the charging, compressing and exhausting of the working fluid. In contrast, the turbine engine eliminates the three ’idle’ strokes, thus enabling more fuel to be burnt in a shorter time;hence it produces a greater power output for a given size of engine.

Rolls-Royce/Snecma Olympus

Rolls-Royce RB37 Derwent 1

A straight-through version of the reverse-flow Power Jets W2B, known as the W2B/26, was developed by the Rover Company from 1941 to 1943. Taken over by Rolls-Royce in April 1943 and renamed the Derwent, it passed a 100hr. test at 2000 lb thrust in November 1943 and was flown at that rating in April 1944. The engine powered the Gloster Meteor III which entered service in 1945.

A turbo-rocket engine.

Fig. 1-12 A turbo-rocket engine.

20. The turbo-rocket engine (fig. 1-12) could be considered as an alternative engine to the turbo/ram jet; however, it has one major difference in that it carries its own oxygen to provide combustion,
21. The engine has a low pressure compressor driven by a multi-stage turbine; the power to drive the turbine is derived from combustion of kerosine and liquid oxygen in a rocket-type combustion chamber. Since the gas temperature will be in the order of 3,500 deg. C, additional fuel is sprayed into the combustion chamber for cooling purposes before the gas enters the turbine. This fuel-rich mixture (gas) is then diluted with air from the compressor and the surplus fuel burnt in a conventional afterburning system.
22. Although the engine is smaller and lighter than the turbo/ram jet, it has a higher fuel consumption. This tends to make it more suitable for an interceptor or space-launcher type of aircraft that requires high speed, high altitude performance and normally has a flight plan that is entirely accelerative and of short duration.

A turbo/ram jet engine.

Fig. 1-10 Comparative propulsive efficiencies

16. At aircraft speeds below approximately 450 miles per hour, the pure jet engine is less efficient than a propeller-type engine, since its propulsive efficiency depends largely on its forward speed; the pure turbo-jet engine is, therefore, most suitable for high forward speeds. The propeller efficiency does, however, decrease rapidly above 350 miles per hour due to the disturbance of the airflow caused by the high blade-tip speeds of the propeller. These characteristics have led to some departure from the use of pure turbo-jet propulsion where aircraft operate at medium speeds by the introduction of a combination of propeller and gas turbine engine.
17. The advantages of the propeller/turbine combination have to some extent been offset by the introduction of the by-pass, ducted fan and propfan engines. These engines deal with larger comparative airflows and lower jet velocities than the pure jet engine, thus giving a propulsive efficiency (Part 21) which is comparable to that of the turbo-prop and exceeds that of the pure jet engine (fig. 1-10).

Mechanical arrangement of gas turbine engines.

Fig. 1-9-2

Basic mechanics

 

Fig. 1-9-1

Mechanical arrangement of gas turbine engines.

PRINCIPLES OF JET PROPULSION

Fig. 1-4 Hero’s engine - probably the earliest
form of jet reaction.

6. Jet propulsion is a practical application of Sir Isaac Newton’s third law of motion which states that, ’for every force acting on a body there is an opposite and equal reaction’. For aircraft propulsion, the ’body’ is atmospheric air that is caused to accelerate as it passes through the engine. The force required to give this acceleration has an equal effect in the opposite direction acting on the apparatus producing the acceleration. A jet engine produces thrust in a similar way to the engine/propeller combination. Both propel the aircraft by thrusting a large weight of air backwards (fig. 1-3), one in the form of a large air slipstream at comparatively low speed and the other in the form of a jet of gas at very high speed.
7. This same principle of reaction occurs in all forms of movement and has been usefully applied in many ways. The earliest known example of jet reaction is that of Hero’s engine (fig. 1-4) produced as a toy in 120 B.C. This toy showed how the momentum of steam issuing from a number of jets could impart an equal and opposite reaction to the jets themselves, thus causing the engine to revolve.

Fig. 1-5 A garden sprinkler rotated by the
reaction of the water jets.

8. The familiar whirling garden sprinkler (fig. 1-5) isa more practical example of this principle, for themechanism rotates by virtue of the reaction to thewater jets. The high pressure jets of modern

Basic mechanics for JET ENGIN

INTRODUCTION
1. The development of the gas turbine engine as an aircraft power plant has been so rapid that it is difficult to appreciate that prior to the 1950s very few people had heard of this method of aircraft propulsion. The possibility of using a reaction jet had interested aircraft designers for a long time, but initially the low speeds of early aircraft and the unsuitably of a piston engine for producing the large high velocity airflow necessary for the ‘jet’ presented many obstacles.

2. A French engineer, René Lorin, patented a jet propulsion engine (fig. 1-1) in 1913, but this was an athodyd (para. 11) and was at that period impossible to manufacture or use, since suitable heat resisting materials had not then been developed and, in the second place, jet propulsion would have been extremely inefficient at the low speeds of the aircraft of those days. However, today the modern ram jet is very similar to Lorin’s conception.

3. In 1930 Frank Whittle was granted his first patent for using a gas turbine to produce a propulsive jet, but it was eleven years before his engine completed its first flight. The Whittle engine formed the basis of the modern gas turbine engine, and from it was developed the Rolls-Royce Welland, Derwent, Nene and Dart engines. The Derwent and Nene turbo-jet engines had world-wide military applications; the Dart turbo-propeller engine became world famous as the power plant for the Vickers Viscount aircraft. Although other aircraft may be fitted; with later engines termed twin-spool, triple-spool, by-pass, ducted fan, unducted fan and propfan, these are inevitable developments of Whittle’s early engine.

Rolls-Royce RB183 Mk 555

Rolls-Royce RB183 Mk 555

Rolls-Royce B23 Welland

Rolls-Royce B23 Welland

On 1 April, 1943, Rolls-Royce assumed responsibility for the Power Jets W2B which, a month earlier, had made its first flight in the Gloster E28/39 at 1200lb thrust. Later known as the B23 Welland it was, during April, put through a 100 hr test at the design rating of 1600 Ib thrust. In June, 1943, it flew in a Gloster Meteor at 1400lb thrust. Production Welland-Meteors were in action against V-1 flying bombs in August 1944.

Rolls-Royce Trent 800

Developed from the RB211, the Trent covers a thrust range of 71,000 lb to 92,000 lb thrust, with the capability to grow beyond 100,000 lb. The Trent 800 features a 110 inch diameter wide-chord fan, high flow compressors and Full Authority Digital Engine Control (FADEC).
Detailed engineering design began in 1988 to meet the propulsion requirements of the Airbus A330 (Trent 700) and Boeing 777 (Trent 800). The Trent first ran in August 1990, and in January 1994 a Trent 800 demonstrated a world record thrust of 106,087 lb. The engine entered service in March 1995 in the Airbus A330.