HYDRAULIC FLUID PROJECTS

HYDRAULIC  FLUID

 

                 Hydraulic  system  liquids  are  used  primarily  to  transmit  and  distribute  forces  to  various  units  to  be  actuated. Liquids  are  able  to  do  this  because  they  are  almost  incompressible. Pascal’s  Law  states  that  pressure  applied  to  any  part  of  a  confined  liquid  is  transmitted  with  undiminished  intensity  to  every  other  part. Thus  if  a  number  of  passages  exist  in  a  system  pressure  can  be  distributed  through  all  of  them  by  means  of  the  liquid.

                  Manufactures  of  hydraulic  devices  usually  specify  the  type  of  liquid  best  suited  for  use  with  their  equipment, in  view  of  the  working  conditions, the  service  required, temperature  expected  inside  and  outside  the  systems, pressure  the  liquid  must  withstand, the  possibilities  of  corrosion  and  other  conditions  that  must  be  considered. Some  of  the  properties  and  characteristics  that  must  be  considered  when  selecting  a  satisfactory  liquid  for  a  particular  system  are  discussed  below. 

AIRCRAFT HYDRAULIC SYSTEMS PROJECTS

AIRCRAFT  HYDRAULIC  SYSTEMS

                    The  word  hydraulic  is  based  on  the  Greek  word  for  water, and  originally  meant  the  study  of  the  physical  behavior  of  water  rest  and  in  motion. Today  the  meaning has  been  expanded  to  include  the  physical  behavior  of  all  liquids, including  hydraulic  fluid.

 

                     It  has  a  great  role  in  aviation. Early aircrafts  has  hydraulic  brake  system. As  aircraft  became  more  sophisticated  newer  systems  with  hydraulic  power  were  developed.

  

                     Landing  gear, wing  flaps, speed  and  wheel  brake, and  flight  control  surfaces  are  the  different  parts  commonly  operated  by  hydraulic  systems.

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.

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. 

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.

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.

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.

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

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.