ISRO’s 100th mission PSLV-C21 PHOTOS

The Indian Space Research Organisation (ISRO) on Friday said it was all set to launch the historic PSLV-C21 on Sunday morning.

The wholly commercial launch will be the space agency's 100th mission in 49 years. So far it has built 62 satellites and flown 37 launch vehicles.

Prime Minister Manmohan Singh is slated to arrive at the space port of Sriharikota, some 120 km from Chennai and located in coastal Andhra Pradesh, on Saturday evening and witness the launch on 9.51 a.m. on Sunday, ISRO officials said.

The PSLV will carry France’s SPOT-6 earth observation satellite as the primary payload and PROITERES, a small spacecraft built by a team of Osaka Institute of Technology in Japan, as a secondary rider. They will be put into their respective pole-to-pole orbits at a distance of 655 km from Earth.

Scientists at the launch site, Satish Dhawan Space Centre at Sriharikota, began a 51-hour countdown at 6.51 a.m. on Friday, a release said.

The Launch Authorisation Board met on Thursday and cleared the event.

During the run-up, teams associated with the launch will complete filling liquid propellants in the second and fourth stages (PS2 and PS4) of the launch vehicle. The rocket and the spacecraft will be checked. Batteries will be charged and the fuel tanks on the satellites will be pressurised, the ISRO officials said.

“Readiness of various ground systems such as tracking radar systems and communication networks will also be ascertained,” they said.

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Aeronautical definitions

Wing geometry 
The planform of a wing is the shape of the wing seen on a plan view of the aircraft. Figure 1.4 illustrates this and includes the names of symbols of the various para- meters of the planform geometry. Note that the root ends of the leading and trailing edges have been connected across the fuselage by straight lines. An alternative to this
convention is that the leading and trailing edges, if straight, are produced to the aircraft centre-line.

Manufacture - Inspection

INSPECTION

Fig. 22-15 Advanced integrated manufacturing system (A.I.M.S.).

47. During the process of manufacture, component parts need to be inspected to ensure defect free engines are produced. Using automated machinery and automated inspection, dimensional accuracy is maintained by using multi-directional applied probes that record sizes and position of features.

Manufacture - machining (E.D.M.) Composite materials and sandwich casings

COMPOSITE MATERIALS AND SANDWICH CASINGS 

Fig. 22-14 Some composite material applications.

43. High power to weight ratio and low component costs are very important considerations in the design of any aircraft gas turbine engine, but when the function of such an engine is to support a vertical take-off aircraft during transition, or as an auxiliary power unit, then the power to weight ratio becomes extremely critical.

Manufacture - Electro-discharge

ELECTRO-DISCHARGE MACHINING (E.D.M.) 
38. This type of machining removes metal from th workpiece by converting the kinetic energy of electric sparks into heat as the sparks strike the workpiece.

Fig. 22-13 Electro-discharge machining

circuit

39. An electric spark results when an electric potential between two conducting surfaces reaches the point at which the accumulation of electrons has acquired sufficient energy to bridge the gap between the two surfaces and complete the circuit. At this point, electrons break through the dielectric medium between the conducting surfaces and, moving from negative (the tool electrode) to positive (the workpiece), strike the latter surface with great energy; fig, 22-13 illustrates a typical spark erosion circuit.
40. When the sparks strike the workpiece, the heat is so intense that the metal to be removed is instan- taneously vaporized with explosive results. Away from the actual centre of the explosion, the metal is torn into fragments which may themselves be melted by the intense heat. The dielectric medium, usually paraffin oil. pumped into the gap between the tool electrode and the workpiece, has the tendency to quench the explosion and to sweep away metallic vapour and molten particles. 

Manufacture - Capillary drilling

Capillary drilling

Fig. 22-12 Typical automated manufacture of compressor blades.

36. Similar in process to stem drilling but using tubes produced from glass incorporating a core of platinum wire (cathode). A twenty per cent nitric acid solution is passed through the tube onto the workpiece and is capable of producing holes as small as 0.009 in. diameter. Depth of the hole is up to forty times greater than the tube in use and therefore determined by tube diameter.
37. Automation has also been added to the process of electro-chemical machining (E.C.M.) with the intro- duction of 360 degree E.G. machining of small compressor blades, ref. fig. 22-12. For some blades of shorter length airfoil, this technique is more cost effective than the finished shaped airfoil when using precision forging techniques. Blades produced by E.C.M. employ integrated vertical broaching machines which take pre-cut lengths of bar material, produce the blade root feature, such as a fir-tree, and then by using this as the location, fully E.C.M. from both sides to produce the thin airfoil section in one operation.

Manufacture - Stem drilling

Stem drilling
35. This process consists of tubes (cathode) produced from titanium and suitably insulated to ensure a reaction at the tip. A twenty per cent solution of nitric acid is fed under pressure onto the blade producing holes generally in the region of 0.026 in. diameter. The process is more speedy in operation than electro-discharge machining and is capable of drilling holes up to a depth two hundred times the diameter of the tube in use.

Manufacture - machining (E.C.M.)

33. Electrolytic grinding employs a conductive wheel impregnated with abrasive particles. The wheel is rotated close to the surface of the workpiece, in such a way that the actual metal removal is achieved by electro-chemical means. The by-products, which would inhibit the process, are removed by the sharp particles embodied in the wheel.
34. Stem drilling and capillary drilling techniques are used principally in the drilling of small holes, usually cooling holes, such as required when producing turbine blades.

Manufacture - Electro-chemical

ELECTRO-CHEMICAL MACHINING (E.C.M.)

Fig. 22-11 Electro-chemical machining.

30. This type of machining employs both electrical and chemical effects in the removal of metal. Chemical forming, electro-chemical drilling and elec- trolytic grinding are techniques of electro-chemical machining employed in the production of gas turbine engine components.

31. The principle of the process is that when a current flows between the electrodes immersed in a solution of salts, chemical reactions occur in which metallic ions are transported from one electrode to another (fig. 22-11). Faraday's law of electrolysis explains that the amount of chemical reaction produced by a current is proportional to the quantityof electricity passed.
32. In chemical forming, (fig. 22-11), the tool electrode (the cathode) and the workpiece (the anode) are connected into a direct current circuit. Electrolytic solution passes, under pressure, through the tool electrode and metal is removed from the work gap by electrolytic action. A hydraulic ram advances the tool electrodes into the workpiece to form the desired passage.

Manufacture - Electron beam welding (E.B.W.)

Electron beam welding (E.B.W.)

Fig. 22-9 Electron beam welding.

Fig. 22-10 Examples of T.I.G. and E.B. welds.

29. This system, which can use either low or high voltage, uses a high power density beam ofelectrons to join a wide range of different materials and of varying thickness. The welding machine ref. fig. 22-9, comprises an electron gun, optical viewing system, work chamber and handling equipment, vacuum pumping system, high or low voltage power supply and operating controls. Many major rotating assemblies for gas turbine engines are manufac- tured as single items in steel, titanium and nickel alloys and joined together i.e., intermediate and high pressure compressor drums. This technique allows design flexibility in that distortion and shrinkage are reduced and dissimilar materials, to serve quite different functions, can be homogeneously joined together. For example, the H.P. turbine stub shafts requiring a stable bearing steel welded to a material which can expand with the mating turbine disc. Automation has been enhanced by the application of computer numerical control (C.N.C.) to the work handling and manipulation. Seam tracking to ensure that the joint is accurately followed and close loop under bead control to guarantee the full depth of material thickness is welded. Focus of the beam is controlled by digital voltmeters. See fig. 22-10 for weld examples.

Manufacture - Tungsten inert gas (T.I.G.) welding

Tungsten inert gas (T.I.G.) welding 

Fig. 22-7 Typical tungsten inert gas 

welding details

Fig. 22-8 Tungsten inert gas welding.

28. The most common form of tungsten inert gas welding, fig, 22-7, in use is the direct current straight polarity i.e., electrode negative pole. This is widely used and the most economical method of producing high quality welds for the range of high strength/high temperature materials used in gas turbine engines. For this class of work, high purity argon shielding gas is fed to both sides of the weld and the welding torch nozzle is fitted with a gas lens to ensure maximum efficiency for shielding gas coverage. A consumable four per cent thoriated tungsten electrode, together with a suitable non-contact method o! arc starting is used and the weld current is reduced in a controlled manner at the end of each weld to prevent the formation of finishing cracks. All welds are visually and penetrant inspected and in addition, weld associated with rotating parts i.e., compressor and/or turbine are radiologically examined to Quality Acceptance Standards. During welding operations and to aid in the control of distortion and shrinkage the use of an expanding fixture is recommended and, whenever possible, mechanised welding employed together with the pulsed arc technique is preferred. A typical T.I.G. welding operation is illustrated in fig. 22-8.

Manufacture - Welding

WELDING
27. Welding processes are used extensively in the fabrication of gas turbine engine components i.e., resistance welding by spot and seam, tungsten inert gas and electron beam are amongst the most widely used today. Care has to be taken to limit the distortion and shrinkage associated with these techniques.

Manufacture - Forging

FORGING

Fig. 22-2 Precision forging.

15. The engine drive shafts, compressor discs, turbine discs and gear trains are forged to as near optimum shape as is practicable commensurate with non-destructive testing i.e., ultrasonic, magnetic particle and penetrant inspection. With turbine and compressor blades, the accurately produced thin airfoil sections with varying degrees of camber and twist, in a variety of alloys, entails a high standard of precision forging, ret. fig. 22-2. Nevertheless precision forging of these blades is a recognised practice and enables one to be produced from a shaped die with the minimum of further work.

16. The high operating temperatures at which the turbine discs must operate necessitates the use of nickel base alloys. The compressor discs at the rear end of the compressor are produced from creep- resisting steels, or even nickel base alloys, because of the

Manufacture - Fabrication

FABRICATION

Fig. 22-6 Wide chord fan blade

construction

25. Major components of the gas turbine engine i.e. bearing housings, combustion and turbine casings, exhaust units, jet pipes, by-pass mixer units and low pressure compressor casings can be produced as fabricated assemblies using sheet materials such as stainless steel titanium and varying types of nickel alloys.

26. Other fabrication techniques for the manufacture of the low pressure compressor wide chord fan blade comprise rolled titanium side panels assembled in dies, hot twisted in a furnace and finally hot creep formed to achieve the necessary configu- ration. Chemical milling is used to recess the centre of each panel which sandwiches a honeycomb core, both panels and the honeycomb are finally joined together using automated furnaces where an activated diffusion bonding process takes place, ref. fig. 22-6.

Manufacture - Casting

CASTING

Fig. 22-3 Method of producing an engine component by sand casting.

Fig. 22-4 Automatic investment casting.

20. An increasing percentage of the gas turbine engine is produced from cast components using sand casting, ref. fig. 22-3, die casting and investment casting techniques; the latter becoming the foremost in use because of its capability to produce components with surfaces that require no further machining. It is essential that all castings are defect free by the disciplines of cleanliness during the casting process otherwise they could cause component failure.

21. All casting techniques depend upon care with methods of inspection such as correct chemica composition, test of mechanical properties, radiolog- ical and microscopic examination, tensile strength and creep tests. 22. The complexity of configurations together with accurate tolerances in size and surface finish is totally dependent upon close liaison with design, manufacturing, metallurgist, chemist, die maker, furnace operator and final casting.

Manufacture - Manufacturing strategy

MANUFACTURING STRATEGY 
9. Manufacturing is changing and will continue to change to meet the increasing demands of aeroengine components for fuel efficiency, cost and weight reductions and being able to process the materials required to meet these demands.
10. With the advent of micro-processors and extending the use of the computer, full automation of components considered for in house manufacture are implemented in line with supply groups manufac- turing strategy, all other components being resourced within the world-wide supplier network.
11. This automation is already applied in the manufacture of cast turbine blades with the seven cell and computer numerical controlled (C.N.C.) grinding centres, laser hardfacing and film cooling hole drilling by electro-discharge machining (E.D.M.). Families of turbine and compressor discs are produced in flexible manufacturing cells, employing automated guided vehicles delivering palletized components from computerized storage to C.N.C. machining cells that all use batch of one techniques. The smaller blades, with very thin airfoil sections, are produced by integrated broaching and 360 degree electrochemical machining (E.C.M.) while inspection and processing are being automated using the computer.

12. Tolerances between design and manufacturing are much closer when the design specification is matched by the manufacturing proven capability.

Manufacture - Introduction

INTRODUCTION 
1. During the design stages of the aircraft gas turbine engine, close liaison is maintained between design, manufacturing, development and product support to ensure that the final design is a match between the engineering specification and the man- ufacturing process capability. 

Fig. 22-1 Arrangements of a triple-spool turbo-jet engine.

2. The functioning of this type of engine, with its high power-to-weight ratio, demands the highest possible performance from each component. Consistent with this requirement, each component must be manufactured at the lowest possible weight and cost and also provide mechanical integrity through a long service life. Consequently, the methods used during manufacture are diverse and are usually determined by the duties each component has to fulfil.
3. No manufacturing technique or process that In any way offers an advantage is ignored and most available engineering methods and processes are employed in the manufacture of these engines, In some instances, the technique or process may appear by some standards to be elaborate, time consuming and expensive, but is only adopted after confirmation that it does produce maximized  component lives comparable with rig test achieve- ments.

Rolls-Royce RB168 Mk202/Mk203 | Rolls-Royce RB39 Clyde

Rolls-Royce RB168 Mk202/Mk203

Rolls-Royce RB39 Clyde

Encouraged by results obtained from the Trent, Rolls-Royce decided to go ahead with an engine designed from the start as a turbo- prop. Named the Clyde it utilized the axial compressor from the Metrovick F2 as first stage and a scaled up supercharger impeller from a Merlin as second stage. First running in August 1945 at 2000 shp, later engines produced up to 4200 shp.

FUEL CONSUMPTION AND POWER-TO-WEIGHT RELATIONSHIP

FUEL CONSUMPTION AND POWER-TO-WEIGHT RELATIONSHIP
46. Primary engine design considerations, particu- larly for commercial transport duty, are those of low specific fuel consumption and weight. Considerable improvement has been achieved by use of the by- pass principle, and by advanced mechanical and aerodynamic features, and the use of improved materials. With the trend towards higher by-pass ratios, in the range of 15:1, the triple-spool and contra-rotating rear fan engines allow the pressure and by-pass ratios to be achieved with short rotors, using fewer compressor stages, resulting in a lighter and more compact engine.

Fig. 21-10 International Standard Atmosphere.

47. S.f.c. is directly related to the thermal and propulsive efficiencies; that is, the overall efficiency of the engine. Theoretically, high thermal efficiency requires high pressures which in practice also means high turbine entry temperatures. In a pure turbo-jet engine this high temperature would result in a high jet velocity and consequently lower the propulsive efficiency (para. 40). However, by using the by-pass principle, high thermal and propulsive efficiencies can be effectively combined by bypassing a proportion of the L.P. compressor or fan delivery air to lower the mean jet temperature and velocity as referred to in para. 43. With advanced technology engines of high by-pass and overall pressure ratios, a further pronounced improvement in s.f.c. is obtained.
48. The turbines of pure jet engines are heavy because they deal with the total airflow, whereas the turbines of by-pass engines deal only with part of the flow; thus the H.P. compressor, combustion chambers and turbines, can be scaled down. The increased power per lb. of air at the turbines, to take advantage of their full capacity, is obtained by the increase in pressure ratio and turbine entry temperature. It is clear that the by-pass engine is lighter, because not only has the diameter of the high pressure rotating assemblies been reduced but the engine is shorter for a given power output. With a low by-pass ratio engine, the weight reduction compared with a pure jet engine is in the order of 20 per cent for the same air mass flow.

PROPULSIVE EFFICIENCY

PROPULSIVE EFFICIENCY

37. Performance of the jet engine is not only concerned with the thrust produced, but also with the efficient conversion of the heat energy of the fuel into kinetic energy, as represented by the jet velocity, and the best use of this velocity to propel the aircraft forward, i.e. the efficiency of the propulsive system.
38. The efficiency of conversion of fuel energy to kinetic energy is termed thermal or internal efficiency and, like all heat engines, is controlled by the cycle pressure ratio and combustion temperature. Unfortunately, this temperature is limited by the thermal and mechanical stresses that can be tolerated by the turbine. The development of new materials and techniques to minimize these limitations is continually being pursued.
39. The efficiency of conversion of kinetic energy to propulsive work is termed the propulsive or external efficiency and this is affected by the amount of kinetic energy wasted by the propelling mechanism. Waste energy dissipated in the jet wake, which represents a is the waste velocity. It is therefore apparent that at the aircraft lower speed range the pure jet stream wastes considerably more energy than a propeller system and consequently is less efficient over this range.
However, this factor changes as aircraft speed increases, because although the jet stream continues to issue at a high velocity from the engine its velocity relative to the surrounding atmosphere is reduced and, in consequence, the waste energy loss is reduced. reference to fig. 21-9 it can be seen that for aircraft designed to operate at sea level speeds below approximately 400 m.p.h. it is more effective to absorb the power developed in the jet engine by gearing it to a propeller instead of using it directly in the form of a pure jet stream. The disadvantage of the propeller at the higher aircraft speeds is its rapid fall off in efficiency, due to shock waves created around the propeller as the blade tip speed approaches Mach 1.0. Advanced propeller technology, however, has produced a multi-bladed, swept back design capable of turning with tip speeds in excess of Mach 1.0 without loss of propeller efficiency. By using this design of propeller in a contra-rotating configuration, thereby reducing swirl losses, a 'prop-fan' engine, with very good propulsive efficiency capable of operating efficiently at aircraft speeds in excess of 500 m.p.h. at sea level, can be produced.

Effect of temperature

Effect of temperature

Fig. 21-7 The effect of altitude on s.h.p. and

fuel consumption

33. On a cold day the density of the air increases so that the mass of air entering the compressor for a given engine speed is greater, hence the thrust or s.h.p, is higher. The denser air does, however, increase the power required to drive the compressor or compressors; thus the engine will require more fuel to maintain the same engine speed or will run at a reduced engine speed if no increase in fuel is available.
34. On a hot day the density of the air decreases, thus reducing the mass of air entering the compressor and, consequently, the thrust of the engine for a given r.p.m. Because less power will be required to drive the compressor, the fuel control system reduces the fuel flow to maintain a constant engine rotational speed or turbine entry temperature, as appropriate; however, because of the decrease in air density, the thrust will be lower. At a temperature of 45 deg.C., depending on the type of engine, athrust loss of up to 20 per cent may be experienced. This means that some sort of thrust augmentation, such as water injection (Part 17), may be required.

Effect of altitude

Effect of altitude

Fig. 21-6 The effects of altitude on thrust

and fuel consumption

32. With increasing altitude the ambient air pressure and temperature are reduced. This affects the engine in two interrelated ways: The fall of pressure reduces the air density and hence the mass airflow into the engine for a given engine speed. This causes the thrust or s.h.p. to fall. The fuel control system, as described in Part 10, adjusts the fuel pump output to match the reduced mass airflow, so maintaining a constant engine speed. The fall in air temperature increases the density of the air, so that the mass of air entering the compressor for a given engine speed is greater. This causes the mass airflow to reduce at a lower rate and so compensates to some extent for the loss of thrust due to the fall in atmospheric pressure. At altitudes above 36,089 feet and up to 65,617 feet, however, the temperature remains constant, and the thrust or s.h.p. is affected by pressure only. Graphs showing the typical effect of altitude on thrust, s.h.p, and fuel consumption are illustrated in fig. 21-6 and fig. 21-7.

ENGINE THRUST IN FLIGHT

ENGINE THRUST IN FLIGHT

Fig. 21-3  Thrust recovery with aircraft

speed.

17. Since reference will be made to gross thrust, momentum drag and net thrust, it will be helpful to define these terms: from Part 20, gross or total thrust is the product of the mass of air passing through the engine and the jet velocity at the propelling nozzle, expressed as: The net thrust or resultant force acting on the aircraft in flight is the difference between the gross thrust and the momentum drag.
18. From the definitions and formulae stated in para, 17; under flight conditions, the net thrust of the Effect of forward speed

19. Since reference will be made to 'ram ratio' and Mach number, these terms are defined as follows: Ram ratio is the ratio of the total air pressure at the engine compressor entry to the static air pressure at the air intake entry. Mach number is an additional means of measuring speed and is defined as the ratio of the speed of a body to the local speed of sound. Mach 1.0 therefore represents a speed equal to the local speed of sound.

Fig. 21-4 The effect of aircraft speed onthrust and fuel consumption.

20. From the thrust equation in para. 18, it is apparent that if the jet velocity remains constant, independent of aircraft speed, then as the aircraft speed increases the thrust would decrease in direct proportion. However, due to the 'ram ratio' effect from the aircraft forward speed, extra air is taken into the engine so that the mass airflow and also the jet velocity increase with aircraft speed. The effect of this tends to offset the extra intake momentum drag due to the forward speed so that the resultant loss of net thrust is partially recovered as the aircraft speed increases. A typical curve illustrating this point is shown in fig. 21-3. Obviously, the 'ram ratio' effect, or the return obtained in terms of pressure rise at entry to the compressor in exchange for the unavoidable intake drag, is of considerable importance to the turbo-jet engine, especially at high speeds. Above speeds of Mach 1.0, as a result of the formation ofshock waves at the air intake, this rate of pressure rise will rapidly decrease unless a suitably designed air intake is provided (Part 23); an efficient air intake is necessary to obtain maximum benefit from the ram ratio effect.

ENGINE THRUST ON THE TEST BENCH

ENGINE THRUST ON THE TEST BENCH

8. The thrust of the turbo-jet engine on the test bench differs somewhat from that during flight. Modern test facilities are available to simulate atmospheric conditions at high altitudes thus providing a means of assessing some of the performance capability of a turbo-jet engine in flight without the engine ever leaving the ground. This is important as the changes in ambient temperatur and pressure encountered at high altitudes consider-ably influence the thrust of the engine.

9. Considering the formula derived in Part 20 for engines operating under 'choked' nozzle conditions, it can be seen that the thrust can be further affected by a change in the mass flow rate of air through the engine and by a change in jet velocity. An increase in mass airflow may be obtained by using wate injection (Part 17) and increases in jet velocity by using afterburning (Part 16).
10. As previously mentioned, changes in ambien pressure and temperature considerably influence the thrust of the engine. This is because of the way they affect the air density and hence the mass of ai entering the engine for a given engine rotationa speed. To enable the performance of similar engines to be compared when operating under differen climatic conditions, or at different altitudes, correction factors must be applied to the calculations to return the observed values to those which would be found under I.S.A. conditions. For example, the thrus  correction for a turbo-jet engine is: Thrust (lb.) (corrected) =

11. The observed performance of the turbo- propeller engine is also corrected to I.S.A. conditions, but due to the rating being in s.h.p. and not in pounds of thrust the factors are different. For example, the correction for s.h.p. is: S.h.p. (corrected) = In practice there is always a certain amount of jet thrust in the total output of the turbo-propeller engine and this must be added to the s.h.p. The correction for jet thrust is the same as that in para. 10.

12. To distinguish between these two aspects of the power output, it is usual to refer to them as s.h.p. and thrust horse-power (t.h.p.). The total equivalent horse-power is denoted by t.e.h.p. (sometimes e.h.p.) and is the s.h.p. plus the s.h.p. equivalent to the net jet thrust. For estimation purposes it is taken that, under sea- level static conditions, one s.h.p. is equivalent to approximately 2.6 lb. of jet thrust. Therefore :

Temperature and pressure notation of a typical turbo-jet engine

INTRODUCTION

Fig. 21-1 Temperature and pressure notation of a typical turbo-jet engine.

1. The performance requirements of an engine are obviously dictated to a large extent by the type of operation for which the engine is designed. The power of the turbo-jet engine is measured in thrust, produced at the propelling nozzle or nozzles, and that of the turbo-propeller engine is measured in shaft horse-power (s.h.p.) produced at the propeller shaft. However, both types are in the main assessed on the amount of thrust or s.h.p. they develop for a given weight, fuel consumption and frontal area.

Rolls-Royce RB168 MK807 | Blackburn Nimbus

Rolls-Royce RB168 MK807

Blackburn Nimbus

 The Nimbus was developed from the A129 turbo-shaft which, in its turn, was a modified Turbomeca Artouste built under licence. The Nimbus developed 968 hp, but for helicopter use was flat-rated at 710 hp. The engine was used in Westland Wasp and Scout helicopters and four 700 hp units were used to power the experimental 5RN-2 hovercraft.

Thrust distribution Inclined combustion chambers & AFTERBURNING

Inclined combustion chambers
22. In the previous example (Para. 14) the flow through the combustion chamber is axial, however, if the combustion chamber is inclined towards the axis of the engine, then the axial thrust will be less than for an axial flow chamber. This thrust can be obtained by multiplying the sum of the outlet thrust by the cosine of the angle (see fig. 20-2). The cosine =  Hypotenuse /Base and for a given angle  is obtained by consulting a table of cosines. It should be emphasized that if the inlet and outlet are at different angles to the engine axis, it is necessary to multiply the inlet and outlet thrusts separately by the cosine of their respective angles.
AFTERBURNING
23. When the engine is fitted with an afterburner (Part 16), the gases passing through the exhaust system are reheated to provide additional thrust. The effect of afterburning is to increase the volume of the exhaust gases, thus producing a higher exit velocity at the propelling nozzle.+

Fig. 20-2 A hypothetical combustion chamber showing values required for calculating thrust

24. Assuming that an afterburner jet pipe and propelling nozzle are fitted to the engine used in the previous

Thrust distribution Engine

Engine
19. It will be of interest to calculate the thrust of the engine by considering the engine as a whole, as the resultant thrust should be equal to the sum of the individual gas loads previously calculated.

Thrust distribution Propelling nozzle

Propelling nozzle

17. The conditions at the inlet to the propelling nozzle are the same as the conditions at the jet pipe outlet, i.e. 16,745 lb.  Therefore, given that the propelling nozzle-- It is emphasized that these are basic calculations and such factors as the effect of air offtakes have been ignored.

18. Based on the individual calculations, the sum of the forward or positive loads is 57,836 lb. and the sum of the rearward or negative loads is 46,678 lb. Thus, the resultant (gross or total) thrust is 11,158 lb.

Thrust distribution Exhaust unit and jet pipe

Exhaust unit and jet pipe
16. The conditions at the inlet to the exhaust unit are the same as the conditions at the turbine outlet, i.e. 14,326 lb.  Therefore, given that the jet pipe--

Thrust distribution Turbine assembly

Turbine assembly 
15. The conditions at the turbine inlet are the same as the conditions at the combustion chamber outlet, i.e. 55,417 lb.

Thrust distribution Combustion chambers

Combustion chambers
14. The conditions at the combustion chamber inlet are the same as the conditions at the diffuser outlet, i.e. 21,235 lb. Therefore, given that the combustion chamber-

Thrust distribution CALCULATING THE THRUST OF THE ENGINE & Compressor casing

CALCULATING THE THRUST OF THE ENGINE
11. When applying the above method to calculate the individual thrust loads on the various components it is assumed that the engine is static. The effect of aircraft forward speed on the engine thrust will be dealt with in Part 21. In the following calculations 'g' is taken to be 32 for convenience. To assist in these calculations the locations concerned are illustrated by a number of small diagrams.

Compressor casing
12. To obtain the thrust on the compressor casing it is necessary to calculate the conditions at the inlet to the compressor and the conditions at the outlet from the compressor. Since the pressure and the velocity at the inlet to the compressor are zero, it is only necessary to consider the force at the outlet from the compressor. Therefore, given that the compressor-

Thrust distribution METHOD OF CALCULATING THE THRUST FORCES

METHOD OF CALCULATING THE THRUST FORCES
7. The thrust forces or gas loads can be calculated for the engine, or for any flow section of the engine, provided that the areas, pressures, velocities and mass flow are known for both the inlet and outlet of the particular flow section.
8. The distribution of thrust forces shown in fig. 20- 1 can be calculated by considering each component in turn and applying some simple calculations. The thrust produced by the engine is mainly the product of the mass of air passing through the engine and the velocity increase imparted to it (i.e. Newtons Second Law of Motion), however, the pressure difference between the inlet to and the outlet from the particular flow section will have an effect on the overall thrust of the engine and must be included in the calculation.
9. To calculate the resultant thrust for a particular flow section it is necessary to calculate the total thrust at both inlet and outlet, the resultant thrust being the difference between the two values obtained.
10. Calculation of the thrust is achieved using the following formula:

Where A = Area of flow section in sq.in.

Thrust distribution DISTRIBUTION OF THE THRUST FORCES

DISTRIBUTION OF THE THRUST FORCES
2. The diagram in fig. 20-1 is of a typical single- spool axial flow turbo-jet engine and illustrates where the main forward and rearward forces act. The origin of these forces is explained by following the engine working cycle shown in Part 2.

Fig. 20-1 Thrust distribution of a typical single-spool axial flow engine.

3. At the start of the cycle, air is induced into the engine and is compressed. The rearward accelera- tions through the compressor stages and the resultant pressure rise produces a large reactive force in a forward direction. On the next stage of its journey the air passes through the diffuser where it  exerts a small reactive force, also in a forward direction,
4. From the diffuser the air passes into the combustion chambers (Part 4) where it is heated, and in the consequent expansion and acceleration of the gas large forward forces are exerted on the chamber walls.

Introduction Thrust distribution

Introduction 
Distribution of the thrust forces 
Method of calculating thrust forces 
Calculating the thrust the engine 

1. Compressor casing
2. Diffuser duct
3. Combustion chambers
4. Turbine assembly
5. Exhaust unit and jet pipe
6. Propelling nozzle
7. Engine
8. Inclined combustion cham

Afterburning 


INTRODUCTION
1. Although the principles of jet propulsion (see Part 1) will be familiar to the reader, the distribution of the thrust forces within the engine may appear somewhat obscure- These forces are in effect gas loads resulting from the pressure and momentum changes of the gas stream reacting on the engine structure and on the rotating components. They are in some locations forward propelling forces and in others opposing or rearward forces. The amount that the sum of the forward forces exceeds the sum of the rearward forces is normally known as the rated thrust of the engine.

Noise suppression CONSTRUCTION AND MATERIALS

CONSTRUCTION AND MATERIALS
17. The corrugated or lobe-type noise suppressor forms the exhaust propelling nozzle and is usually a separate assembly bolted to the jet pipe. Provision is usually made to adjust the nozzle area so that it can be accurately calibrated. Guide vanes are fitted to the lobe-type suppressor to prevent excessive losses by guiding the exhaust gas smoothly through the lobes to atmosphere. The suppressor is a fabricated welded structure and is manufactured from heat- resistant alloys.
18. Various noise absorbing lining materials are used on jet engines. They fall mainly within two categories, lightweight composite materials that are used in the lower temperature regions and fibrous- metallic materials that are used in the higher temperature regions. The noise absorbing material consists of a perforate metal or composite facing skin, supported by a honeycomb structure on a solid backing skin which is bonded to the parent metal of the duct or casing. For details of manufacture of these materials refer to Part 22.

Noise suppression CONSTRUCTION AND MATERIALS,Introduction Noise suppression, Noise suppression Engine noise, Noise suppression Methods of suppressing noise, Noise suppression Construction and materials, 

Noise suppression Noise absorbing materials and location METHODS OF SUPPRESSING NOISE 1

Fig. 19-6 Noise absorbing materials and location.

13. Deep corrugations, lobes, or multi-tubes, give the largest noise reductions, but the performance penalties incurred limit the depth of the corrugations or lobes and the number of tubes. For instance, to achieve the required nozzle area, the overall diameter of the suppressor may have to be increased by so much that excessive drag and weight results. A compromise which gives a noticeable reduction in noise level with the least sacrifice of engine thrust, fuel consumption or addition of weight is therefore the designer's aim.

Noise suppression METHODS OF SUPPRESSING NOISE ypes of noise suppressor

METHODS OF SUPPRESSING NOISE

Fig. 19-5 Types of noise suppressor.

10. Noise suppression of internal sources is approached in two ways; by basic design to minimize noise originating within or propagating from the engine, and by the use of acoustically absorbent linings. Noise can be minimized by reducing airflow disruption which causes turbulence. This is achieved by using minimal rotational and airflow velocities and reducing the wake intensity by appropriate spacing between the blades and vanes. The ratio between the number of rotating blades and stationary vanes can also be advantageously employed to contain noise within the engine.


11. As previously described, the major source of noise on the pure jet engine and low by-pass engine is the exhaust jet, and this can be reduced by inducing a rapid or shorter mixing region. This reduces the low frequency noise but may increase the high frequency level. Fortunately, high frequencies are quickly absorbed in the atmosphere and some of the noise which does propagate to the listener is beyond the audible range, thus giving the perception of a quieter engine. This is achieved by increasing the contact area of the atmosphere with the exhaust gas stream by using a propelling nozzle incorporating a corrugated or lobe-type noise suppressor (fig. 19-5).

Noise suppression ENGINE NOISE

Fig. 19-4 Comparative noise sources of low and high by-pass engines.

6. Compressor and turbine noise results from the interaction of pressure fields and turbulent wakes from rotating blades and stationary vanes, and can be defined as two distinct types of noise; discrete tone (single frequency) and broadband (a wide range of frequencies). Discrete tones are produced by the regular passage of blade wakes over the stages downstream causing a series of tones and harmonics from each stage. The wake intensity is largely dependent upon the distance between the rows of blades and vanes. If the distance is short then there is an intense pressure field interaction which results in a strong tone being generated. With the high bypass engine, the low pressure

Rolls-Royce RM60 | Rolls-Royce Conway

Rolls-Royce Conway

Rolls-Royce RM60

Produced in response to an Admiralty contract for a coastal-craft engine with good cruising economy, the RM60, although based on aeroengine philosophy, was designed from the first as a marine gas turbine. Two RM60s went to sea in 1953 in the former steam gunboat HMS Grey Goose, the world's first warship to be powered solely by gas turbines.