GB2179702A - Turbo-propeller aircraft engine - Google Patents
Turbo-propeller aircraft engine Download PDFInfo
- Publication number
- GB2179702A GB2179702A GB08521418A GB8521418A GB2179702A GB 2179702 A GB2179702 A GB 2179702A GB 08521418 A GB08521418 A GB 08521418A GB 8521418 A GB8521418 A GB 8521418A GB 2179702 A GB2179702 A GB 2179702A
- Authority
- GB
- United Kingdom
- Prior art keywords
- exhaust
- propeller
- exit
- gas turbine
- turbine engine
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Withdrawn
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- 239000007789 gas Substances 0.000 claims abstract description 159
- 239000000446 fuel Substances 0.000 claims abstract description 36
- 238000011144 upstream manufacturing Methods 0.000 claims description 18
- 230000000063 preceeding effect Effects 0.000 claims 1
- 238000000034 method Methods 0.000 abstract description 11
- 238000002485 combustion reaction Methods 0.000 description 5
- 230000015572 biosynthetic process Effects 0.000 description 1
- 230000035939 shock Effects 0.000 description 1
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
- B64D33/04—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of exhaust outlets or jet pipes
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C25/00—Alighting gear
- B64C25/32—Alighting gear characterised by elements which contact the ground or similar surface
- B64C25/42—Arrangement or adaptation of brakes
- B64C25/423—Braking devices acting by reaction of gaseous medium
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/46—Nozzles having means for adding air to the jet or for augmenting the mixing region between the jet and the ambient air, e.g. for silencing
- F02K1/50—Deflecting outwardly a portion of the jet by retractable scoop-like baffles
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Aviation & Aerospace Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Supercharger (AREA)
Abstract
The present invention relates to turbo-propeller aircraft gas turbine engines of the pusher type, that is engines which have the propellers arranged at the downstream end of the engine and normally in the flow of exhaust gases. Overfuelling of pusher turboprops during starting gives a wet start, in which burning fuel and exhaust gases impinge upon and cause damage to the propeller blading. To prevent this damage during starting of the turbo-propeller engine, the engine is provided with diverter means, to direct at least a portion of the exhaust gases and burning fuel away from the propellers. The embodiment in Fig. 3 shows a diverter means (44) comprising blocker doors (52, 54) and a translating cowl (50) which move between a first position in which the doors (52, 54) block an exhaust duct (32) and direct the exhaust gases and burning fuel through a second exit (58) away from the propellers (28, 30) during starting procedure, and a second position in which the doors (52, 54) and cowl (50) block the second exit (58) and the exhaust gases are directed through the propeller blades 28. Various alternative diverter means are disclosed. <IMAGE>
Description
SPECIFICATION
Improvements in or relating to turbo-propeller aircraft engines
The present invention relates to aircraft engines of the turbo-propeller type, commonly known as turboprops.
These engines usually have a gas generator comprising a compressor, a combustion system, and a compressor driving turbine, a power turbine driven by the exhaust gases from the compressor driving turbine and arranged to drive one or more propellers through reduction gearing.
The propeller or propellers are usually arranged at the upstream end of the engine which means that the air entering the gas generator has to pass through the propeller blading, before entering an annular or chintype engine air intake. Because of the obstruction of the propeller blading and the tortuous flow path to the gas generator, the inlet conditions are not ideal.
The solution to the intake problem is to position the propeller or propellers downstream of the gas generator so that there is no obstruction to the gas generator air intake. However, this makes it necessary to discharge the hot, high velocity engine exhaust gases either through the propeller blading, or through the roots of the propeller blades, as in our copending published application GB 2117054A.
A- further possibility is to duct the exhaust gases away from the engine and to position the propeller or propellers out of the way of the exhaust gases as in UK patent no.
809811 and our co-pending published application GB 2138507A.
Turbo-props of the type with the propeller or propellers positioned at the downstream end of the gas generator are generally known as pusher turbo-props. Pusher turbo-props suffer from a problem in that during starting of the engine they are sometimes overfueled and this gives a "wet start", when burning fuel and exhaust gases can impinge upon the propeller blading causing damage.
The present invention seeks to provide a turbopropeller aircraft engine of the pusher type which prevents burning fuel and exhaust gases impinging upon the propeller during engine starting procedures.
Accordingly the present invention provides a turbopropeller aircraft gas turbine engine of the pusher type comprising in flow series a gas generator comprising an unobstructed air intake, compressor means, combustion means, and a compressor driving turbine means, a power turbine driven by the exhaust gases from the compressor driving turbine, the power turbine driving a reduction gear box, and at !east one propeller driven from the reduction gear box, an exhaust duct having an exhaust exit upstream of the at least one propeller for directing the exhaust gases in a downstream direction through the at least one propeller, diverter means to direct at least a portion of the exhaust gases and burning fuel away from the at least one propeller during starting of the gas turbine engine to prevent damage to the at least one propeller.
The exhaust duct may have a second exhaust exit upstream of the exhaust exit, the diverter means comprising door means moveable between a first position in which the door means blocks the exhaust exit and directs substantially all the exhaust gases and burning fuel through the second exhaust exit away from the at least one propeller during starting of the gas turbine engine to prevent damage to the at least one propeller and a second position in which the door means blocks the second exhaust exit and the exhaust gases are directed through the exhaust exit during normal operation of the gas turbine engine.
The exhaust duct may be annular and have an annular exhaust exit, the diverter means comprising blocker doors hinged to a nacelle and moveable between said first position in which the blocker doors extend across and block the annular exhaust duct and annular exhaust exit downstream of the second exhaust exit to direct the exhaust gases and burning fuel through the second exhaust exit away from the at least one propeller during starting of the gas turbine engine, and said second position in which the blocker doors move across and block the second exhaust exit and the exhaust gases are directed through the annular exhaust exit during normal operation of the gas turbine engine.
The diverter means may also comprise a translating cowl axially moveable between a first position in which the translating cowl is positioned axially downstream of the second exhaust exit to allow the exhaust gases and burning fuel to flow through the second exhaust exit and a second position in which the translating cowl blocks the second exhaust exit.
The translating cowl and blocker doors may move along a plurality of screwjacks.
The exhaust exit may be of the chin-type, the diverter means comprising a clamshell door rotatable between said first position in which the clamshell door extends across and blocks the exhaust duct and chin-type exhaust exit at the downstream end of the second exhaust exit to direct the exhaust gases and burning fuel through the second exhaust exit away from the at least one propellor during starting of the gas turbine engine, and said second position in which the clamshell door moves across and blocks the second exhaust exit and the exhaust gases are directed through the chin-type exhaust exit during normal operation of the gas turbine engine.
The exhaust duct may be annular and have an annular exhaust exit, the diverter means comprising blocker doors hinged to an inner annular wall of the annular exhaust duct and moveable between said first position in which the blocker doors extend across and block the annular exhaust duct and annular exit downstream of the second exhaust exit to direct the exhaust gases and burning fuel through the second exhaust exit away from the at least one propeller during starting of the gas turbine engine, and said second position in which the blocker doors move to open the annular exhaust duct and annular exhaust exit to direct the exhaust gases through the annular exhaust exit during normal operation of the gas turbine engine.
The area of the second exhaust exit may be larger than the area of the exhaust exit in order to reduce back pressure in the gas turbine engine during starting to reduce the possibility of a wet start.
The exhaust duct may have a second exhaust exit upstream of the exhaust exit, the exhaust duct comprising a strut extending radially from the gas turbine engine upstream of the at least one propeller, the exhaust exit comprising one or more exits in the strut for directing exhaust gases in a downstream direction over the surface of the strut, the diverter means comprising a valve positioned within said second exhaust exit and rotatable between a first position in which the valve directs a portion of the exhaust gases and burning fuel through the second exhaust exit away from the at least one propeller during starting of the gas turbine engine to increase the exit area to reduce the possibility of a wet start occurring and a second position in which the valve blocks the second exhaust exit and the exhaust gases are directed through the exhaust exit and over the surface of the strut during normal operation of the gas turbine engine.
The diverter means may comprise door means positioned axially between the exhaust exit and the at least one propeller, the door means being hinged to a nacelle and moveable between a first position in which the door means extends radially from the nacelle to direct substantially all the exhaust gases and burning fuel leaving the exhaust exit away from the at least one propeller during starting of the gas turbine engine to prevent damage to the at least one propeller and a second position in which the door means extends axially to allow the exhaust gases to flow through the propeller during normal operation of the gas turbine engine.
The diverter means may be moved between the first and second positions when the high pressure air in the compressor reaches a predetermined value.
The diverter means may be moved between the first and second positions pneumatically.
The engine may have a pair of contra-rotating propellers, and the propeller or propellers may be of the propfan type.
The present invention will now be more particularly described with reference to the accompanying drawings in which:Figure 1 shows a turbo-propeller aircraft gas turbine engine of the pusher type having diverter means according to the present invention.
Figure 2 shows a turbo-propeller aircraft gas turbine engine of the pusher type having an alternative diverter means according to the present invention.
Figure 3 is a sectional view to a larger scale of the diverter mean shown in figure 1.
Figure 4 is a sectional view to a larger scale of the diverter means shown in figure 2.
Figure 5 shows a turbo-propeller aircraft gas turbine engine of the pusher type having a further diverter means according to the present invention.
Figure 6 is a sectional view to a larger scale of the diverter means shown in figure 5.
Figure 7 is a sectional view to a larger scale of an alternative embodiment of a diverter means shown in figure 1.
and figure 8 shows a turbo-propeller aircraft gas turbine engine of the pusher type having a further diverter means according to the present invention.
Referring to figures 1 and 2, a turbo-propeller aircraft gas turbine 10 of the pusher type comprises a gas generator 12 having an air inlet 14, a compressor 16, a combustion system 18 and a compressor driving turbine 20.
The exhaust gases from the gas generator 12 drive a power turbine 22 which in turn drives a pair of contra-rotating multi-bladed propellers 28, 30 via a reduction gear box 26 and shaft 24. In this arrangement the gas generator 12, the power turbine 22, shaft 24, gear box 26 and propellers 28, 30 have a common axis of rotation.
The exhaust gases leaving the power turbine flow into an annular duct 32 and are conveyed to an annular exit 34. The annular exit 34 is positioned upstream of the propellers 28, 30 which are located downstream of the gas generator 12 of the gas turbine engine
10. Therefore it can be seen that the exhaust gases from the gas generator flow in a downstream direction along arrows A, as shown in the top half of figure 3, along duct 32 through the annular exit 34 onto and through the pm- pellers 28, 30, during normal operation of the engine.
This arrangement suffers from a problem during starting of the gas turbine engine 10, as mentioned previously, because the combustion system 18 can be overfueled and not all of the fuel will be burnt in the combustion system. The exhaust gases from the gas generator 12 will therefore contain burning fuel, which is directed onto the propellers, causing damage.
The invention provides a diverter means 44 which prevents the burning fuel and exhaust gases impinging upon the propellers during engine starting procedures. The diverter means 44 comprises a translating cowl 50, and rings of blocker doors 52, 54 and screwjacks 56.
The blocker doors 52, 54 are hinged together, and the doors 54 are hinged at their opposite end to the nacelle 46 of the engine 10. The nacelle 46 has an exit 58 upstream of the annular exit 34, and the blocker doors 54 are hinged to the nacelle 46 at the downstream end of the exit 58. The doors 52 are hinged to nuts 60 which are free to move on the screwjacks 56. The screwjacks 56 extend axially across the exit 58.
When the gas turbine engine is at rest the diverter means 44 is arranged so that the blocker doors 52 and 54 extend across and block the annular duct 32 downstream of the exit 58 and the translating cowl 50 is at a position downstream of the exit 58 so that the exit 58 is open, as shown in the bottom half of figure 3.
During the starting procedure, any burning fuel in the exhaust gases flowing through the annular duct 32, is directed with the exhaust gases to flow through the exit 58 and away from the propellers by the blocker doors 52 which extend across the annular duct 32, to prevent the exhaust gases and burning fuel flowing through annular exit 34 and onto the propellers.
The area of the exit 58 can be chosen to be greater than that of annular exit 34, this aids in starting the gas turbine engine 10 as there is a reduced back pressure, and this reduces the possibility of a wet start occurring.
Once the engine is running the diverter means 44 takes up the position shown in the top half of figure 3. This is achieved by operation of the screwjacks 56 which cause the nuts 60 to move in an upstream direction along the screwjacks 56, the translating cowl 50 moves axially upstream across and closes the exit 58, and the blocker doors 52 and 54 rotate about their hinges and move across and close the exit 58 while opening the annular duct 32 for the flow of exhaust gases to and through the annular exit 34.
The screwjacks 56 can be operated automatically when the pressure in the compressor reaches a predetermined value to open the annular exit 34, and when the pressure falls below this value, i.e. on engine shut down the diverter means 44 blocks the annular duct 32.
The high pressure air from the compressor 16 can be used to pneumatically operate the screwjacks 56 and diverter means 44 such a system is well known in the art and will not be described further.
The exit 58 could be fully annular, but may be in the form of one or more arcs of an annulus so as not to direct the burning fuel into undesired areas, ie towards the aircraft fuslage or wings.
Referring to figures 2 and 4, which show a turbopropellor aircraft gas turbine engine 10 of the pusher type similar to that in figures 1 and 3. The gas generator 12 is identical to that in figures 1 and 3, and like parts are denoted by like numerals. The power turbine 22 drives a pair of contra-rotating propellers of the propfan type, ie of thin section and highly swept to delay shock wave formation and having the ability to operate at a high power loading. The propellers 36, 38 are driven by the power turbine via a reduction gear box 26 and shafts 24. In this arrangement the propeller axis of rotation is offset from the axis of rotation of the gas generator 12 and power turbine 22.
The exhaust gases leaving the power turbine flow into a duct 40 and are conveyed to a chin-type exit 42. The chin exit 42 is positioned upstream of the propellers 36, 38 which are located downstream of the gas generator 12 of the gas turbine engine 10. The exhaust gases flow in a downstream direction along arrows C, as shown in figure 4, along duct 40 and through chin exit 42 into and through the propellers 36, 38 during normal operation of the engine.
This arrangement has a diverter means 66 which prevents the burning fuel and exhaust gases impinging upon the propellers during engine starting procedures. The diverter means 66 comprises a clamshell door 70 which is arranged to rotate about a pivot 74 on the engine nacelle 68. The nacelle 68 has an exit 72 upstream of the chin exit 42.
When the gas turbine engine is at rest the diverter means 66 is arranged so that the clamshell door 70 extends across and blocks the duct 40 and the exit 72 is open, as shown in phantom in figure 4. During the starting procedure, any burning fuel, in the exhaust gases flowing through the duct 40 is directed with the exhaust gases to flow through the exit 72 and away from the propellers as shown by arrow D, by the clamshell door 70 which extends across the duct 40 to prevent the exhaust gases and burning fuel flowing through the chin exit 42 and onto the propellers.
Again the area of exit 72 can be chosen to be greater than the chin exit 42 so as to aid in the starting procedure by reducing the back pressure.
After the engine has started the diverter means takes up the position shown in figure 4. This is achieved by rotation of the clamshell door 70 about the pivots 74 from the position where the clamshell door blocks the duct 40 to a position where the clamshell door blocks the opening 72 and allows the exhaust gases to flow along duct 40 and through the chin exit 42.
Figure 5 and 6 show a turbopropeller aircraft gas turbine engine 10 of the pusher type similar to that in figures 1 and 3. The gas generator, 12 is indentical to that in figures 1 and 3, and like parts are denoted by like numerals. The power turbine 22 drives a pair of contra-rotating propellers 28, 30 via a reduction gear box 26 and shaft 24. In this arrangement the propeller axis of rotation is coaxial with axis of rotation of the gas generator 12 and power turbine 22.
The exhaust gases leaving the power turbine flow into a duct 80 within the nacelle 92, and are conveyed to a duct 84 formed within a strut 82. The strut 82 is positioned upstream of the propellers and downstream of the power turbine and extends radially from the gas turbine engine 10 and may form part of a structure to carry the engine from an aircraft.
The duct 84 has a nozzle 86 which may be used to direct the exhaust gases in a downstream direction over the tail of the aircraft, as described more fully in our co-pending published application GB 2138507A. The strut 82 can be provided with exits 88, generally slots with vanes, to discharge exhaust gases in a downstream direction over the strut to suppress the wake. The exhaust gases flowing through the openings 88 in a downstream direction over the strut 82 are depicted by arrows E in figure 6.
This arrangement has a diverter means 90 which prevents a portion of the burning fuel and exhaust gases impinging upon the propellers during engine starting procedures. The diverter means 90 comprises a valve 94 which is arranged to rotate about a pivot 98 on the engine nacelle 92. The nacelle 92 has an exit 96 upstream of the duct 84 in which the valve 94 is positioned.
When the gas turbine engine is at rest the diverter means 90 is arranged so that the exit 96 is open, as shown in phantom in figure 6.
During the starting procedure, a portion of the burning fuel and exhaust gases flowing through duct 80 is directed to flow through the exit 96 and away from the propellers, as shown by arrow F. The ducts 80 and 84 are not blocked, but by the opening of the valve 94 the total exit area for the exhaust gases is greater than that for exits 86 and 88 alone.
This aids in the starting procedure by reducing back pressure and minimising the possibility of a wet start occurring.
Once the engine has started the valve 94 takes up the position shown in figure 6. This is achieved by rotation of the valve 94 about the pivot 98, to close the exit 96.
The exit 96 and valve 94 in the nacelle 92 will generally be positioned at the bottom of the nacelle 92, not opposite the strut 82 as is shown in figure 6 for convenience. This is so that the unburnt fuel deposited on the nacelle may run out of the duct 80.
Referring to figure 7 which shows diverter means 100 applied to a turbo-propeller aircraft gas turbine engine similar to that in figure 3.
The diverter means 100 comprises a ring of blocker doors 102 secured to an inner annular wall 48 of the annular exhaust duct 32 by hinges 104. The blocker doors 102 are rotatable about the hinges 104, and in a first position they extend across and close the annular duct 32 at the downstream end of the second annular exit 58 to direct the exhaust gases and burning fuel along arrow G through the exit 58 during the engine starting procedure. The blocker doors 102 rotate about their hinges 104 once the engine has started to allow the exhaust gases to flow through the annular exit 34, as shown by arrow G. A translating cowl 50 or other suitable means must be provided to close off the second exit 58 during normal operation of the engine, but which moves to open the exit 58 when the engine has stopped.
Figure 8 shows a diverter means 110 applied to a pusher type turbo-propeller aircraft gas turbine engine 10 similar to that in figure 1. The diverter means 110 comprises a ring of door means 114 positioned axially between the exhaust exit 34 of the exhaust duct 32 and the at least one propeller 28, 30. The door means 114 is hinged to a nacelle 112 and is rotatable about the hinges from a first position in which the door means 114 extend radially from the nacelle to direct the exhaust gases and burning fuel leaving the exhaust exit 34 away from the propellers 28, 30 to a second position in which the door means 114 extends axially flush with the nacelle 112 to allow the exhaust gases to flow through the propellers 28, 30.
The embodiments in figures 2, 4, 5, 6, 7 and 8 can also be automatically operated by high pressure air from the compressor 16.
The high pressure air from the compressor pneumatically operating the clamshell door valve or doors when the pressure reaches a predetermined value to close off the respec tive second exit.
The invention is not restricted to the type of gas generator, reduction gear box, propeller or number of propellers. In the embodiments shown two propellers have been shown, but a single propeller could be driven by the power turbine. The propeller could be of the normal type or of the propfan type.
The embodiments described, and shown, have an exhaust exit upstream of the propeller or propellers for directing the exhaust gases in a downstream direction through the propellers.
A diverter means is arranged to direct a portion of the exhaust gases and burning fuel away from the propeller or propellers during starting of the gas turbine engine to reduce damage to the propeller or propellers.
Claims (15)
1. A turbo-propeller aircraft gas turbine engine of the pusher type comprising in flow series a gas generator comprising an unobstructed air intake, compressor means, a power turbine driven by the exhaust gases from the compressor driving turbine, the power turbine driving a reduction gear box, and at least one propeller driven from the reduction gear box, an exhaust duct having an exhaust exit upstream of the at least one propeller for directing the exhaust gases in a downstream direction through the at least one propeller, diverter means to direct at least a portion of the exhaust gases and burning fuel away from the least one propeller during starting of the gas turbine engine to prevent damage to the at least one propeller.
2. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in claim 1 in which the exhaust duct has a second exhaust exit upstream of the exhaust exit, the diverter means comprising door means moveable between a first position in which the door means blocks the exhaust exit and directs substantially all the exhaust gases and burning fuel through the second exhaust exit away from the at least one propeller during starting of the gas turbine engine to prevent damage to the at least one propeller and a second position in which the door means blocks the second exhaust exit and the exhaust gases are directed through the exhaust exit during normal operation of the gas turbine engine.
3. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in claim 2 in which the exhaust duct is annular and has an annular exhaust exit, the diverter means comprising blocker doors hinged to a nacelle and moveable between said first position in which the blocker doors extend across and block the annular exhaust duct and annular exhaust exit downstream of the second exhaust exit to direct the exhaust gases and burning fuel through the second exhaust exit away from the at least one propeller during starting of the gas turbine engine, and said second position in which the blocker doors move across and block the second exhaust exit and the exhaust gases are directed through the annular exhaust exit during normal operation of the gas turbine engine.
4. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in claim 3 in which the diverter means comprises a translating cowl axially moveable between a first position in which the translating cowl is positioned axially downstream of the second exhaust exit to allow the exhaust gases and burning fuel to flow through the second exhaust exit and a second position in which the translating cowl blocks the second exhaust exit.
5. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in claim 4 in which the translating cowl and blocker doors move along a plurality of screwjacks.
6. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in claim 2 in which the exhaust exit is of the chin type, the diverter means comprising a clamshell door rotatable between said first position in which the clamshell door extends across and blocks the exhaust duct and chin type exhaust exit at the downstream end of the second exhaust exit to direct the exhaust gases and burning fuel through the second exhaust exit away from the at least one propeller during starting of the gas turbine engine, and said second position in which the clamshell door moves across and blocks the second exhaust exit and the exhaust gases are directed through the chin type exhaust exit during normal operation of the gas turbine engine.
7. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in claim 2 in which the exhaust duct is annular and has an annular exhaust exit, the diverter means comprising blocker doors hinged to the inner annular wall of the annular exhaust duct and moveable between said first position in which the blocker doors extend across and block the annular exhaust and annular exhaust exit downstream of the second exhaust exit to direct the exhaust gases and burning fuel through the second exhaust exit away from the at least one propeller during starting of the gas turbine engine, and said second position in which the blocker doors move to open the annular exhaust duct and annular exhaust exit to direct the exhaust gases through the annular exhaust exit during normal operation of the gas turbine engine.
8. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in any of claims 2 to 7 in which the area of the second exhaust exit is larger than the area of the exhaust exit in order to reduce back pressure in the gas turbine engine during starting to reduce the possibility of a wet start.
9. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in claim 1 in which the exhaust duct has a second exhaust exit upstream of the exhaust exit, the exhaust duct comprising a strut extending radially from the gas turbine engine upstream of the at least one propeller, the exhaust exit comprising one or more exits in the strut for directing exhaust gases in a downstream direction over the surface of the strut, the diverter means comprising a valve positioned within said second exhaust exit and rotatable between a first position in which the valve directs a portion of the exhaust gases and burning fuel through the second exhaust exit away from the at least one propeller during starting of the gas turbine engine to increase the exit area to reduce the possibility of a wet start occurring and a second position in which the valve blocks the second exhaust exit and the exhaust gases are directed through the exhaust exit and over the surface of the strut during normal operation of the gas turbine engine.
10. A turbo-propeller aircraft gas turbine en gine of the pusher type as claimed in claim 1 in which the diverter means comprises door means positioned axially between the exhaust exit and the at least one propeller, the door means being hinged to a nacelle and moveable between a first positon in which the door means extends radially from the nacelle to direct substantially all the exhaust gases and burning fuel leaving the exhaust exit away from the at least one propeller during starting of the gas turbine engine to prevent damage to the at least one propeller and a second position in which the door means extends axially to allow the exhaust gases to flow through the propeller during normal operation of the gas turbine engine.
11. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in any of claims 2 to 10 in which the diverter means is moved between the first and second positions when the high pressure air in the compressor reaches a predetermined value.
12. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in claim Ilin which the diverter means is moved between the first and second positions pneumatically.
13. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in any of claims 1 to 12 in which the engine has a pair of contra-rotating propellers.
14. A turbo-propeller aircraft gas turbine engine of the pusher type as claimed in any of the preceeding claims in which the propeller or propellers are of the propfan type.
15. A turbo-propeller aircraft gas turbine engine of the pusher type substantially as herein described with reference to and as illustrated in figures 1 and 3, or figures 2 and 4 or figures 5 and 6.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB08521418A GB2179702A (en) | 1985-08-28 | 1985-08-28 | Turbo-propeller aircraft engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB08521418A GB2179702A (en) | 1985-08-28 | 1985-08-28 | Turbo-propeller aircraft engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| GB2179702A true GB2179702A (en) | 1987-03-11 |
Family
ID=10584390
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB08521418A Withdrawn GB2179702A (en) | 1985-08-28 | 1985-08-28 | Turbo-propeller aircraft engine |
Country Status (1)
| Country | Link |
|---|---|
| GB (1) | GB2179702A (en) |
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4964844A (en) * | 1987-09-05 | 1990-10-23 | Rolls-Royce Plc | Gearbox arrangement for driving coaxial contra rotating multi-bladed rotors |
| GB2430013A (en) * | 2005-09-09 | 2007-03-14 | Christian Koenig | Propeller drive |
| CN103029826A (en) * | 2012-12-10 | 2013-04-10 | 江西洪都航空工业集团有限责任公司 | Aircraft heat protection and electric energy extraction integrated structure |
| CN113062801A (en) * | 2021-04-19 | 2021-07-02 | 中国航发湖南动力机械研究所 | Power rear output type turboprop engine and airplane |
| EP3984889A1 (en) * | 2020-10-14 | 2022-04-20 | Pratt & Whitney Canada Corp. | Aircraft propulsion system with propeller and cooling fan |
Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2138507A (en) * | 1983-04-22 | 1984-10-24 | Rolls Royce | Mounting end exhausting in turbo-propellor aircraft engines |
-
1985
- 1985-08-28 GB GB08521418A patent/GB2179702A/en not_active Withdrawn
Patent Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2138507A (en) * | 1983-04-22 | 1984-10-24 | Rolls Royce | Mounting end exhausting in turbo-propellor aircraft engines |
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4964844A (en) * | 1987-09-05 | 1990-10-23 | Rolls-Royce Plc | Gearbox arrangement for driving coaxial contra rotating multi-bladed rotors |
| GB2430013A (en) * | 2005-09-09 | 2007-03-14 | Christian Koenig | Propeller drive |
| US7886544B2 (en) * | 2005-09-09 | 2011-02-15 | Christian Koenig | Propeller or propeller drive |
| CN103029826A (en) * | 2012-12-10 | 2013-04-10 | 江西洪都航空工业集团有限责任公司 | Aircraft heat protection and electric energy extraction integrated structure |
| CN103029826B (en) * | 2012-12-10 | 2016-04-20 | 江西洪都航空工业集团有限责任公司 | Aircraft thermal protection and power extraction integral structure |
| EP3984889A1 (en) * | 2020-10-14 | 2022-04-20 | Pratt & Whitney Canada Corp. | Aircraft propulsion system with propeller and cooling fan |
| US11719248B2 (en) | 2020-10-14 | 2023-08-08 | Pratt & Whitney Canada Corp. | Aircraft propulsion system with propeller and cooling fan |
| CN113062801A (en) * | 2021-04-19 | 2021-07-02 | 中国航发湖南动力机械研究所 | Power rear output type turboprop engine and airplane |
| CN113062801B (en) * | 2021-04-19 | 2022-07-22 | 中国航发湖南动力机械研究所 | Power rear output type turboprop engine and airplane |
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