User:Pieter1961/sandbox: Difference between revisions – Wikipedia

Finished writing a draft article? Are you ready to request review of it by an experienced editor for possible inclusion in Wikipedia? Submit your draft for review!

Engine support structure, accessories ref Young book FRames and structures (GE4 design part 1)

Intake responsibility of airframe manufacturer Cumpsty book p.209

Lubrication requirements, spool bearings, AGB, RGB (PW1100 construction) (GEnX inflight shutdown)
AGB roller shaft bearings (CFM LEAP p.101) (bearings seals secondary air) sump pressurisation (CFM56-5C)

secondary air (ASME secondary air)

Thrust to aircraft, mounts (PW1100 construction) (Trent 1000) (LEAP)

 Treager Fig.15-20

This section gives intuitive explanations for where pressure rise comes from, why a turbine is needed, why pressure rise is limited unless variable features are added to compressor, what causes compressor to stop working properly (surge).

i) Fast -spinning blades grab air which acquires some proportion of the blade speed (whirl velocity/angular momentum increase). The air now has kinetic energy, part of which can be converted to pressure energy using a suitably shaped passage to catch the air (diffuser).

With the momentum increase comes an equal and opposite force on the blades. (The straight-line momentum increase inside a jet engine has an equal and opposite force (thrust) pushing the aircraft forward). The force on the blades tending to slow them is balanced by the turbine blades keeping them going.

Note: turbine examples are chosen according to availability of relevant photographs.

Turbines used in jet engines take energy out of hot, high pressure gas to drive the engine compressors, and the pumps for the engine oil and fuel systems. They are known as gas turbines, with similar-looking blades to steam turbines.

A turbine stage consists of one row of stationary vanes followed by a row of rotor blades. Stages are added together to give a turbine enough power to drive its compressor. The example turbine from an early turbojet, the General Electric J79, needed 3 stages to drive its 17-stage compressor.

Jet engines usually have two turbines, high pressure and low pressure. The high pressure turbine (HPT) is part of the so-called hot section of the engine (the other part is the combustor). The hot section is exposed to higher operating temperatures than the rest of the engine and deteriorates more rapidly so needs inspecting sooner.^[1][2]

The low pressure turbine (LPT) is in the much cooler (and lower pressure) gas leaving the HPT. Most of the HPT power is fed back into the engine (to the compressor) to raise the air pressure going into the combustor. The LPT drives the fan at the front of the engine.

There is a noticeable similarity in layout between the turbine and an axial compressor. They both have stages of stationary vanes and fast-moving blades.

• a row of stationary vanes followed by a row of moving blades (constitutes one turbine stage)
• rotating disc holding blades
• shaft connecting with compressor
• turbine roller bearing, located in a frame to support turbine shaft radially. Frame bolted to turbine cases to form engine structural backbone^[3][4]
• casing with small blade tip clearance to save fuel
• tip clearance control system which changes casing temperature to prevent rubs while changing thrust settings and to maintain small clearance when steady running (cruise)

Source:^[5]

The accompanying picture of an early turbojet turbine shows some of these parts (an exhaust cone would hide the disc when fitted in an aircraft). There is a row of blades attached to the rim of the turbine disc. The blades run in a close-fitting casing with as small a clearance as can be achieved without rubbing. By the time the gas from the combustor leaves the blades (approaching the viewer) it has removed thousands of horse power (which equates to lowered pressure and temperature) and fed it back to the front of the engine to raise the ambient pressure for efficient combustion. It passes the viewer at a pressure greater than ambient on its way out of the exhaust nozzle. The pressure is an indicator of the thrust of the engine (higher pressure more thrust).^[6]

• 
  This rear view of a single stage turbine (early turbojet Klimov VK-1) shows some basic features. There is a disc with blades attached around the rim. A connecting shaft goes forward to the compressor. There is a close-fitting casing which leaves a small gap at the end of the blades. The blades are hotter than the disc so to limit the effects of heat conduction from the blades cooling air is passed across the disc faces.^[7]
• 

  Shows the Polish equivalent of the Klimov VK-1 single-stage turbine vane ring (the first part of a turbine to be subjected to the combustor gas) with bladed rotor removed. When the rotor is in place the blade tips run inside the silver surface behind the vane ring. The diameter of the silver surface sets the blade tip clearance and the amount of gas which makes no contribution to turning the turbine.

Air leaves the combustor going straight backwards (axially). To give up energy to the moving part of the turbine it needs to be going sideways (whirling) to push the turbine blades round. It also needs to be speeded up. Turning and accelerating the air is done by the first part of the turbine which is a ring of stationary passages (known as nozzle guide vanes or vane ring). The air leaves the nozzles and hits the moving blades at the best angle and speed to transfer energy to the rotor. The moving blades take energy from the air which shows as a drop in pressure and temperature and the energy appears in the form of work available at the other end of the turbine shaft (the compressor).

• Tip clearance closure (rub/pinch point) potentially on every flight.
• Tip clearance reduction long-term (creep)
• Turbine operating temperature required to get take-off thrust increases over time (is an engine health indicator) (distortion and corrosion of hot parts, seal clearances opening up (after rubs), compressor and turbine running clearances opening up (after casing distortion with thrust and manoeuvre loads)
• Turbine vane area back-pressures the compressor (controls pressure ratio, air flow and fuel consumption)^[9][10]
• Turbine speed is held back (compromises turbine design – needs more stages) so as to keep fan speed low enough for good efficiency
• Significance of EGT – representative of TIT^[11]

When the engine is running at high speed and turbine temperature (the two go together with high fuel flow) the blades permanently lengthen (creep) and oxidation and hot corrosion gradually cause the blades to corrode and crack.

• rotor balancing ^[12]
• operating temperatures – airliner, combat aircraft.

When the engine is changing speed (RPM), up or down, the temperature of the blades and disc gradually change towards the steady value which goes with the new turbine temperature. Changing speeds and temperatures cause the blade and disc materials to suffer from metal fatigue (cracks appear and grow in length).

Increasing speed/temperature from the engine idle values for take-off causes the disc to grow and the tip clearances to get smaller, and at the same time the casing heats up and the clearances get bigger. The disc and casing respond to gas temperature changes at different rates (the disc is thick/heavy, the casing is thin/light) and the blades may rub while the metal temperatures are evening out at a steady gas temperature.

The velocity diffusion required on the blade surface increases as the turning between the blades increases.^[13]
In comparison with this compressor flow that through the turbine is following its natural direction to a lower pressure (ultimately ambient outside the engine) so a large amount of turning is possible which means more energy is extracted per stage than in a compressor.^[14]

The axial turbine is essentially the reverse of the axial compressor except for one essential difference: the turbine flow operates under a favorable pressure gradient. This permits greater angular changes, greater pressure changes, greater energy changes, and higher efficiency.^[15]

• Compressor/turbine
• 
  J79 compressor showing blades turning the flow towards the axial direction which results in a diffusing (rising pressure) passage between each blade. Note small amount of turning between each pair of blades which is limited by flow separation when going against the rising pressure. The small amount of turning and energy transfer to the air means a small pressure rise for each stage, a pressure ratio of about 1.17 for this engine.^[16] A 2018 compressor GEnX 27:1 11 stages 2 HPT^[17]
• 

  J79 turbine showing blades turning flow away from axial direction which results in an accelerating passage (nozzle) between each blade. Note the greater amount of turning taking place (compared to compressor blades) which means more energy transfer per row of blades.

Past studies on single stage turbines have shown that the best performance
can be achieved in a configuration with a high rim speed and long blades (high AN2 ) for low throughflow gas velocity.^[18]
For a given output the gas velocities, deflections, and hence losses, are reduced in proportion to the square of higher mean blade speeds^[19]

As well as the force on the blades which cause the turbine disc to turn there is also a force in the axial direction which contributes to the tension in the shaft. Also disc pressure loading ref thrust bearing.
The vane ring is trying to turn in the opposite direction to the blades but no work (energy transfer) is done because it does not move so no pressure and temperature drop.

Radial temperature profile /thermal stresses /fatigue cycles

Main article: Active tip-clearance control
Turbine growths tip clearance ^[20]
Cooling flows schematic ^[21]

cooling flows schematic ^[22]

^[23]

The vanes are internally cooled with air flowing from the compressor. Some of the cooling air also comes out of the visible holes to form a cooling layer on the vane surface. In this way the metal temperature is limited to slow the rate of corrosion (oxidation for example) to give the parts the long life promised by the engine maker.

Purge flow^{[24] [25]}

^[26]

Tip groove ^[27]]]

The compressor flow is turned towards the axial direction which gives a larger exit area than entry (hence diffuser). The turbine flow is turned away from the axial direction giving a smaller exit than entry (hence nozzle).^[28]

compressor stages 3/4x turbine stages ref acc/cecel flow ^[29]
Blade loading diagram is pressure distribution curve, area between the curves represents blade force acting in tangential direction. ^[30]

Turning Angles
For a turbine the inlet pressure to a stage is greater than the exit pressure. Thus, separation will not occur on the blades because of an adverse pressure gradient. As a result, more expansion can be accomplished by a turbine than compression can be accomplished by a compressor. This means that ({Ainlet/Aexitlturbine > {Aexit/Ainlet}compressor). Thus, on the basis of Figure 8.3, the turning angle is greater for a turbine than for a compressor. Because each turbine stage has more flow turning than a compressor stage, more energy is extracted per stage than in a compressor.

Number of Stages
As just noted, more energy is extracted per stage than in a compressor. Since all of the turbine power is used to drive the compressive devices for a jet engine, the number of turbine stages is less than the number of compressor stages.^[31]

Creep and clearance
pheorwm,enon of creep of moving blades, iJo., thPi.r continuous, although slow,
plastic flow. This rlow limits the operation life of moving blades. p.86 Kuz’min

ln procccG of prolon(‘;cd operation, radial clearances in turbine decrease.
flucc to creep of’ material of lhc moving blades and disks, under action of centrifugal
forces trrere oc:cu r·s ti’He:Lr continuous lengthening in ra.ciial direction. In
turbine housings there hac; been observed a reverse phenomenon, i.e,, their
c;hrl.nkae;e and warping. This decrease of clearances usually limits the operation
life of engl.nes. Thus, in the turbine of the VK-1 engine (see Fig. 4.23) initial
rad:Lal clearance betweEm rotor wheel 9 and body 7 is established from ? .3 to 2.6
nun. lts decrease in the proc”ess of operation is allowed up to 1.3 mm, after which
the c:learancc is restored by borinr=; the housing 7.^[32]
Peri.phera.L part of disks
obtainc; inf’lu.x of’ twat from moving blades and directly from working gases, the
-115-
disks have more or less intense cooling by air, and in central part the heat is
ejected into the turbine shaft and other components connected with it. Thus,
temperature of turbine disks decreases from periphery to center. The less
heated central part of disk prevents its more heated peripheral part to be freely
expanded and in the latter there appear circumferential compression stresses
(nev,atlve stresses).p.115 Kuz’min.

Thermal and kinetic energy conversion inside turbine to mechanical in shaft^[33]
Temperature profile GE book Fig 5.26
Book elements of propulsion Mattingly para 9.5
Because a turbine expands from high to low pressure, there is no such thing as turbine surge or stall. The turbine needs fewer stages than the compressor. The blades have more curvature and the gas stream velocities are higher.

Designers must, however, prevent the turbine blades and vanes from melting in a very high temperature and stress environment. Consequently, bleed air extracted from the compression system is often used to cool the turbine blades/vanes internally. Other solutions are improved materials and/or special insulating coatings. The discs must be specially shaped to withstand the high stresses imposed by the rotating blades. Improved materials help to keep disc weight down.

a bare engine is delivered to a podding facility^[34] where the nacelle supplier assembles the inlet and the rest of the nacelle, which includes the thrust reverser, to the engine.^[35]

see ^[36]
“As the air…” para ^[37]
Inlet has to provide uniform airflow to the engine, otherwise the fan will stall. For an airliner inlet it means rounded lip at entrance to inlet. Air not approaching from straight ahead will not be able to hug the surface if lip is sharp. Smooth flow will leave surface and become chaotic, causing fan stall and compressor surging. Most severe for an airliner is take-off when aircraft and its engine pods inclined upwards but moving forwards into the upwash, which combined for angle of attack of 30 degrees.^[38]
Modern supersonic combat aircraft have entry to inlet at side of fuselage where skin boundary layer has thickened and must be prevented from entering inlet. If its not removed it separates with shock waves present at supersonic speeds. Causes compressor surging. It is prevented from entering by spacing inlet away with gap bl thickness or putting a bump in front of inlet. This causes bl to divide and go either side of bump instead of straight over the top and in.

The components already mentioned are those which contribute to producing thrust, because they constitute the solid surfaces which interact with the rearward-flowing gas which flows through all of them in sequence. The momentum and pressure of the gas changes as it flows through each component. There is a force associated with a momentum change (force equals rate of change of momentum) and with a pressure change (a pressure difference acting on its respective area difference is a force).

The components above are linked by a parameter common to all of them, the flow rate of gas passing through the engine which is the same for all components at the same time (as a basic statement this is an acceptable approximation which ignores the addition of fuel in the combustor and bleeding air from the compressor).^[40] There is a common requirement for all of them, to waste as little of the fuel supplied to the engine in collectively contributing to the output of the engine, which is thrust or power to a propeller or rotor. For flow through ducts this means keeping the flow Mach number (Mn) low since losses increase with increasing Mn. Having too high a Mn at entry to a duct is particularly relevant in ducts where there is heat addition, ie the engine combustor, and an afterburner if fitted, since the Mn would reach sonic velocity if the entry Mn were too high (Rayleigh flow).

The compressor and turbine, as well as having to pass the same flow, turn together so the speeds have a fixed relationship (usually equal unless connected with a gearbox), and one drives the other so the turbine power has to equal the compressor power.^[40] At the same time losses in the compressor and turbine have to be reduced so they operate with acceptable efficiency.

The designing, sizing and manipulation of the operating characteristics of the components so they work together as a unit is known as matching.^[41]

By the time the jet engine entered airline service the very powerful and complex, post-war piston engine and propeller combination introduced for crossing oceans was very unreliable. Joe Sutter, Stratocruiser development engineer, said, “The reliability of the P&W R-4360 was a real disappointment. If the engine didn’t fail in flight, very often the propeller would”.^[44] Engine horsepower and the frequency of in-flight failures were directly correlated.^[45]

W.J.Overend, Manager of Performance at Delta Airlines, described early airline experience of gas turbine running time between failures as “phenomenal”, compared to DC-7 piston engines.^[46] The limited reliability of piston engines led to an operating restriction being placed on two engine airplanes 50 years ago. In light of these advances, and because the safety and reliability of two-engine airplanes equal or exceed those of three- or four-engine airplanes, the industry no longer views propulsion system reliability as the primary safety and reliability concern in extended operations. diversion time can be significantly increased without added risk if the IFSD rate is sufficiently low. An IFSD rate of 0.01 per 1,000 engine-hours—or twice the engine reliability level required for 180-min ETOPS has been determined to allow unconstrained operations with two-engine airplanes. ^[47]

A350 ETOPS370^[48]

The jet engine itself has become much more reliable. By the turn of the century high bypass engines were more than 50 times more reliable than the large piston engines of 1950. Advancing technology resulted in 120 minute ETOPS in 1985 which required an in-flight shut down rate of 0.05 for 1,000 engine hours. In 1988 180 required 0.02/1,000 hours, by 2000 rates were 0.01 or 1/2 that required for 180 minute diversions. A variety of aircraft systems and operational issues (cargo fire suppression, facilities at alternate airports)
In the past the focus on flight safety has been on the powerplant, starting with the limited reliability of the piston engine for trans-ocean commercial flights.
^[49]

Regulatory acceptance of increased reliability p209 Borrow Bibel book
IFSD is events per 100,000 hrs .02 per1,000
engine shutdown p212
Engines have become so reliable that regulatory bodies allow twin-engine aircraft to fly routes which are 330 minutes flying time from the nearest airport. This means, if an engine has to be shut down, the aircraft will fly for 5 1/2 hours on one engine as an acceptable practice.

As Andy has already said it comes down to “What is the difference between a fan and a prop?” Gunston has said “the fan can be considered a multibladed shrouded propeller” which might suggest that if you get the blade shape and number right you get a ducted fan. Trying to understand all the important differences in detail (number of blades, spacing between them, their cross section/airfoil shape) can be bypassed by going straight to their respective performance diagrams which are completely different.

It turns out that the propeller has to be turned into a compressor which means the ducted fan isn’t just a different looking set of blades in a duct because that would only be half a compressor. Stationary vanes behind it are the second half. Also the exit area from the duct needs to be right to back-pressure the compressor by the right amount.

The significance of this distinction between prop and fan is that the prop performance depends on aircraft speed because the blades are enveloped in it. Fan performance doesn’t because flight speed doesn’t reach the fan. Prop performance diagrams are constructed with flight speed information. Fan diagrams show things pertinent to a compressor.

The difference is that propeller performance is determined by the speed of the aircraft (and so needs variable angle blades to suit) but a fan is a compressor and its performance is isolated from flight speed. It depends on already-compressed air from the flight speed ram effect, ie the pressure and temperature in front of the blades. All the important/difficult differences in detail (number of blades, spacing between them, their cross section/airfoil shape) can be bypassed by going straight to their respective performance diagrams which are completely different. The fan diagram shows what’s pertinent to a compressor or pump, pressure rise and airflow and effect of outlet area.

The charts which display prop performance all include aircraft speed, because as the propeller screws its way through the air the airflow past the blades is made up from the prop speed and the flight speed. Prop performance charts show forward speed.

What is not obvious looking at the ducted fan is that it’s not just a single set of blades (as with a propeller) spinning in a duct. It’s a compressor and the fan is only the first half. The other half is the stationary vanes behind. Also the duct has a restricted outlet area to back-pressure them to give the required pressure. Fan performance is shown as pressure rise and airflow for different fan speeds and outlet areas.

The prop and the fan, although they are both whirling blades which end up doing the same thing, to cause momentum changes in the approaching air stream (https://arc.aiaa.org/doi/10.2514/6.1981-1566) go about it differently because the requirement from the blades is different. The prop blades are like aircraft wings with the same requirement, lift with minimum drag. The whirling wings are going vertically so so-called lift is thrust and drag opposes the crankshaft or turbine motion, so is the torque part of shaft power. As with wings the angle of attack is important so the blade pitch and aircraft speed have to give a good angle of attack to get the most thrust with the torque/shaft power.

The ducted fan isn’t just a shrouded propeller. It’s a single stage compressor so the spinning blades need a stator row behind and a restricted exit area to build up the required pressure. The way they work is obviously different when looking at their respective performance diagrams. There is no resemblance. What is immediately noticeable is the prop diagrams have flight speed but the fan shows up with no flight speed, but as a compressor showing pressure rise and flow for different speeds. The dependence of propeller performance on flight speed made it obsolete for fighter aircraft from 1945 onwards.ref German jet survey

There are 3 diagrams for the propeller and all include flight speed because the propeller screws its way forwards with a tip speed which increases with aircraft speed. So, how well, or not, it works is inextricably related to aircraft speed, see https://eaglepubs.erau.edu/introductiontoaerospaceflightvehicles/chapter/propellers/ “Advance Ratio, Thrust Coefficient, Power Coefficient, Propulsive Efficiency”

The fan is a compressor so has to spin in a tube with a restricted exit area which builds up some pressure. It raises the air pressure from the level in front of it whether its stationary ambient air or ram-induced from flight. A fan performance diagram, see Figure 5 https://torroja.dmt.upm.es/congresos/asme_2011/data/pdfs/trk-1/GT2011-46397.pdf
doesn’t have flight speed, only fan inlet pressure and temperature, whether flight-derived or not. In other words, for a particular pressure and temperature the performance is the same. Figure 5 shows the fan is operating almost at the same point whether static on the ground, red triangle, or cruise, green circle.

Another angle on the above: when a prop is tested in a wind tunnel the tunnel has to blow air at it at flight speed. A fan, though, is tested in an altitude facility where the air is supplied at the stagnation temp and pressure corresponding to the flight condition. If the intake is to be tested in front of the fan then air is supplied to the aircraft intake at flight Mn and the intake produces the ram conditions for the fan.

Fan blades
Wennerstrom Blades high a/r, could ignore 3D effects . long chord higher pressure before stall and large flow range
high stage pressure ratio Tumansky, 535E4 211 wide chord fan,offspring V2500,

Thrust reverse

Thrust spoiling, 90 deg, only ram drag so OK for high BPR, lower bpr (low ram drag) need to produce rev thr (p.23 Static perf 6 innov)

(p.22 Static perf 6 innov)For cascade type parameters for perf are overall TR effectiveness ie eff rev, fan reverser effectiveness eff fa, and area match beta. O/A 0.31, fan0.54. Transient area match into surge , Aeff less than nozzle.

(NACA 1314) Target type so-called because jet is turned by a target positioned behind exhaust nozzle. 0.5 . Blockers and cascades on early t/j and bypass engines, used Comet 707 C141 typ .4 to 0.5 (Russia book Table 1) Reinjestion Conway HPC failure (Russia p18 ref Flight 1961)

thrust reverser design Embraer para 4 TR requirements, parra 10 Fig 10 CFD image at cancellation speed TR integration with engine and aircraft. Critical condition as reverted jets more intense than free stream FOD re-inj and bouyancy shows isosurface of total temp of plume.

Cold stream pivot door ^[51]

Wheel brakes vg for dry runway, see stopping distances where TR has little extra ^[52], but TR braking independent of surface condition so make big difference to be able to stop on surfaces which have low friction between tires and ground. high speed end of landing run TR most effective when KE greatest ,109158,
powerbacks limited to low thrust tail mounted,,not high bypass due to large amount of airflow 109158. TR essential to achieving max level of operating safety
The necessary TR capability comes from the landing requirement for a particular aircraft.
Max capability for a high bypass might be fan flow reverse and hot nozzle reverse or at least spoiling. with this capability on all engines. Multi engined a/c other than twins, which obviously have both engines equipped, vary in this respect ranging from no engines, BAe146, to 1/3, Dassault Falcons, to 1/2 A380 to all for twin jets. High bypass usually have fan flow reversal only.

The most effective part of the landing run for using reverse thrust is when the aircraft kinetic energy is greatest, as soon as the aircraft is on the ground with highest speed. The engine is at its maximum reverse thrust setting which varies with aircraft type from 100 % of forward take-off thrust to 75%, for example (Fokker 100^[53]). Maximum reverse has to be cancelled to reverse idle thrust before the exhaust intensity overcomes the opposing free stream as the aircraft slows (50 to 70 knots depending on aircraft type). Otherwise the engine will suck in its own exhaust as well as risk foreign object damage.

C17 37 deg fwd cold and hot blocker doors with cascades (Tavernetti C17… )

Known as target type because

Pivot type Reverses all exhaust from a low bypass engine such as

uses blocker doors and turning cascades for high bypass

used with fan duct reverse eg C-17
uses blocker doors and turning cascades for fan air

Happens naturally as a supplement to flow turning devices. Also works with thrust spoiler on engine which doesn’t reverse the flow, just turns through 90 deg

Ram drag is bigger than CF6 fan reverse^[54]

Cite error: There are <ref group=nb> tags on this page, but the references will not show without a {{reflist|group=nb}} template (see the help page).