GB2384827A

GB2384827A - A Fastening Between The C-Duct and Core of a Ducted Fan Gas Turbine Engine.

Info

Publication number: GB2384827A
Application number: GB0202047A
Authority: GB
Inventors: James Martin Townsend
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2002-01-30
Filing date: 2002-01-30
Publication date: 2003-08-06
Also published as: GB0202047D0

Abstract

A ducted fan gas turbine engine has a C-duct assembly having an inner 20 and outer 22 casing which define a nacelle located downstream of a fan case, characterised by an inner casing structure which is releasably attached to the engine core via fastenings 36. The nacelle outer casing may comprise two semi-circular parts attachable to the fan case by an annular projection engaging with a groove in the fan case, to form an annular nacelle each semi-circular part is hingably mounted on the aircraft pylon and has fastenings near and diametrically opposite the pylon support. The nacelle may also include a bifurcation portion 28,30 joining the inner and outer casing, with adjacent fastenings. The fastening between engine and nacelle may include a pin (38 figure 4) and cooperating recess (40 figure 5), which may be mounted on either engine core 16 or nacelle.

Description

Gas Turbine Engines The invention relates to gas turbine engines and particularly to turbo fan engines for use as propulsion units for aircraft.

Gas turbine engines include one or more turbines driven by combustion gases, each turbine in turn driving a fan or compressor via an interconnecting shaft. In a turbo fan engine, a proportion of the air drawn into the engine by the fan bypasses the engine core and provides propulsive thrust.

The bypass air initially flows through an annular duct defined between the engine core and a fan case. Located downstream of the fan case is a "C-duct" nacelle which includes inner and outer casing structures also defining an annular duct for bypass air. Bifurcation walls join the inner and outer casing structures in upper and lower regions of the nacelle. The C-duct nacelle is split into two halves, each of which is hinged at its top to a pylon, which is attached to the body of the aircraft.

The C-duct nacelle can house a thrust reverser. When the thrust reverser is activated, a number of doors each pivot from an inactive position to a position in which they block the flow of bypass air and therefore reverse the thrust of the engine.

A front of the outer casing structure of the nacelle is attached to the fan case via a "V groove" provided on the fan case. As the nacelle is hinged closed, an annular protrusion provided on the nacelle engages the V groove and provides a connection between the outer casing structure of the nacelle and the fan case, around nearly their full circumferential extents.

In most civil aircraft, each engine is mounted on the aircraft via a pylon attached to the engine. When the

hr ; : st reverser (housed in the C duct nacelle) is employed, : : he rearward load generated must be fed into the engine and

thereby through the pylon to the aircraft. In one conventional arrangement, there is no connection of the inner casing structure of the nacelle to the engine. Therefore, the load path from the thrust reverser doors passes to the inner casing structure of the C duct nacelle, through the bifurcation walls to the outer casing structure, from the V groove to the fan case, through the outlet guide vanes which link the fan case with the core and eventually to the core engine and through thrust struts or other thrust transmission devices to the pylon. This is a long and inefficient load path.

Aircraft manoeuvre derived engine loads can bend the aircraft core, leading to detrimental effects on engine performance and reliability. If such bending has to be allowed for, the engine must be designed such that the clearances between the compressor and turbine tips and their associated casings must be relatively large, and this has a detrimental effect on the performance of the engine.

Some engines include an additional inner V groove similar to the V groove described above, but connecting the inner casing structure of the C duct nacelle to the core of the engine. This inner V groove provides advantages in terms of load transfer, but is disadvantageous in that it obstructs the mounting of the pylon on the engine core. In addition, it is necessary to provide structure to support the inner V groove and transmit the loads, and this can prove difficult to achieve.

According to the invention there is provided a turbo fan engine including: an engine core; a fan case surrounding the core and defining a bypass duct between the fan case and the core; a C-duct assembly downstream of the fan case and including inner and outer generally annular casing structures defining a duct for bypass air therebetween; characterised in that the inner casing structure of

the C-duct assembly is releasably attachable to the engine core via discrete fastenings.

Preferably the outer casing structure of the C-duct assembly is attachable to the fan case. The outer casing structure may include an annular projection, which engages an annular groove in the fan case.

Preferably the inner and outer casing structures of the C-duct assembly together form a nacelle which is split into two semi-annular halves, each half being hingably mountable on a pylon of an aircraft, so that the two halves may hinge towards and away from one another. The C-duct assembly is normally mounted generally under the pylon, with its top being hinged thereto.

Preferably the discrete fastenings are provided generally in the circumferential region of the nacelle which is hingably mountable on the pylon. Preferably further discrete fastenings are provided in a region circumferentially opposite to this region.

Preferably the discrete fastenings are provided in a forward, upstream part of the nacelle.

Preferably the nacelle includes a bifurcation joining the inner and outer casing structures generally in the region of the hinge attachment and in the region opposite thereto. In use, the bifurcations are therefore generally in upper and lower regions of the nacelle.

Preferably the discrete fastenings are provided in the regions of the bifurcations. Preferably two discrete fastenings are provided in the region of each bifurcation.

Each discrete fastening may include a pin provided on the engine core or on a member attached to the engine core and a complementary recess provided in the inner casing structure of the nacelle. Alternatively each fastening may include a pin provided on the inner casing structure of the nacelle and a complementary recess provided on or associated with the engine core. The pins are preferably oriented with their axes tangential to the inner and outer

casing structures and therefore parallel to the relative hinging movement between the inner casing structure and the engine core. Each pin may include a chamfered portion and each recess may include a complementary sloping portion. A single elongate member may constitute the pins for each of two fastenings.

An embodiment of the invention will be described for the purpose of illustration only with reference to the accompanying drawings in which: Fig. 1 is a diagrammatic exploded view showing the general arrangement of a known gas turbine engine; Fig. 2A is a part sectional view of a known gas turbine engine illustrating the area in which the fan case connects to the C duct nacelle and Fig. 2B is an end view of Fig. 2A, illustrating the A-frame; Fig. 3 is a highly diagrammatic cross section illustrating the general layout of a gas turbine engine according to the invention; Fig. 4 is a sketch in perspective view of part of the core of a gas turbine engine according to the invention; Fig. 5 is a diagrammatic perspective view of half of the C duct nacelle of a gas turbine engine according to the invention; Fig. 6 is diagrammatic sectional view of the C duct nacelle of a gas turbine engine according to the invention; Fig. 7A and 7B are diagrammatic front views of the C duct nacelle of a gas turbine engine according to the invention, in the closed and opened positions respectively; and Figs. 8 and 9 are diagrammatic sketches of the pin arrangement for attaching the C duct nacelle to the core in a gas turbine engine according to the invention.

Referring to Figs. 1 and 2, a ducted fan gas turbine engine generally indicated at 10 includes an air intake 12 and a propulsive fan 14 housed within fan cowl doors. Downstream of the fan are intermediate and high pressure

compressors, combustion equipment, and high, intermediate and low pressure turbines, located within an engine core 16.

The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 14 to produce two air flows, a first air flow into the intermediate pressure compressor (and subsequently to the high pressure compressor, combustion equipment, and turbines) and second air flow of "bypass air" which provides propulsive thrust.

The compressed air exhausted from the compressors is directed into the combustion equipment where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through and thereby drive the high, intermediate and low pressure turbines before being exhausted to provide additional propulsive thrust. Referring to Fig. 2,-the high, intermediate and low pressure turbines drive the high and intermediate pressure compressors and the fan respectively by suitable interconnecting shafts.

A fan case 18 is attached to the engine core by outlet guide vanes 24 and an "A frame" 25 (see Fig. 2B) which connects the rear of the fan case to the core.

The bypass air flows within a generally annular passageway defined between the fan case 18 and the engine core 16. Behind the fan case 18 there is located a C duct nacelle 19, which includes inner and outer casing structures 20 and 22 respectively (see Figs. 3 and 5 in particular). A top of the C duct nacelle is hinged to a pylon 31 (see Fig. 1) which is attached to an aircraft in use. The nacelle 19 is split down its vertical centre into two semi-annular halves which are able to swing outwardly away from one another (see Fig. 7a and 7b).

The nacelle 19 forms a generally annular passageway for bypass air after it leaves the annular passageway between the fan case 18 and the engine core 16. The

annular passageway in the nacelle 19 is divided at its top and bottom by bifurcation walls 28 and 30 respectively.

The nacelle 19 is connected to the fan case 18 via a V groove 32 (see Fig. 2). A generally annular projection on the outer casing structure 22 of the nacelle 19 engages the annular V groove 32 provided on the fan case 18, when the nacelle 19 is hinged shut. This provides a connection between the nacelle and the fan case nearly all around their circumferential extents. In some engines, an inner V groove is also provided (34 in Fig. 2), this V groove providing a connection between the inner casing structure 20 of the nacelle 19 and the engine core.

The C duct nacelle 19 houses the engine thrust reversers (not illustrated). The thrust reversers include a series of doors which are able to hinge from inactive positions in which they do not affect the flow of air through the nacelle 19 to activated positions in which they block this flow of air and therefore reverse the thrust of the engine.

The engine is mounted on the aircraft via a pylon (not illustrated) attached to the engine core generally in the region of the rear of the fan case. When the thrust reverser (housed in the nacelle 19) is employed, the rearward load generated must be fed into the engine and thereby through the pylon to the aircraft. In one conventional arrangement, there is no connection of the inner casing structure 20 to the engine. Thus, the load path from the thrust reverser doors passes to the inner casing structure 20 of the nacelle 19, through the bifurcation walls 28 and 30 to the outer casing structure 22, from the V groove 32 to the fan case 18 and eventually to the engine core 16 and through thrust struts or other thrust transmission devices to the pylon. This is a long and inefficient load path. Lack of connection between the inner casing structure 20 and the engine core also means that the nacelle 19 offers less resistance to engine

bending due to aircraft manoeuvre.

In particular, the engine may be subject to bending loads as the aircraft takes off. In this case an"intake upload"acts to try to bend the engine core. Some of this load tends to be transmitted via the V groove to the C duct nacelle. If there is only an outer V groove connecting the nacelle 19 with the engine fan case 18, this load is not efficiently shared between the engine and the nacelle.

If such bending has to be allowed for, the engine must be designed such that the clearances between the compressor and turbine tips and their associated casings must be relatively large, this having a detrimental effect on the performance of the engine, and can reduce reliability.

Engines with an additional inner V groove 34 transmit loads more efficiently from the engine core to the C duct nacelle and vice versa. However, the position of the inner V groove has been found to obstruct the mounting of the pylon on the core.

In a gas turbine engine according to the invention, a connection is provided between the inner casing structure 20 of the C duct nacelle 19 and the engine core 16, via discrete fastenings 36. The discrete fastenings 36 take the form of pins 38 provided on the engine core, which engage with recesses 40 provided in the inner casing structure 20 of the nacelle 19. The pins 38 are clearly shown in Fig. 4 and the recesses 40 in Fig. 5. Referring also to Figs. 8 and 9, each pin 38 includes a chamfer 42 which engages a sloping surface 44 of the recess. This ensures that any slight misalignment between the pins 38 and recesses 40 can be accommodated.

Referring to Figs. 7A and 7B, as the C duct nacelle 19 is hinged inwardly from its open position to its closed position, the pins 38 move into engagement with the recesses 40. The pins 38 and recesses 40 are aligned tancentially to the inner casing structure 20, such that rh comte smoothly into engagement on tangential hinging

movement of the nacelle 19.

The discrete fastenings 36 are provided in the regions of the bifurcation walls 28 and 30, in upper and lower front regions of the inner casing structure 20.

The discrete fastenings 36 provide a load path between the engine core 16 and the inner casing structure 20 of the C duct nacelle 19. It has been found that, where an inner V groove is provided, most of the load is transmitted at the top and the bottom of this V groove, in the region of the bifurcation walls 28 and 30. Thus, the discrete fastenings in these regions provide efficient load transfer. When the thrust reversers are employed, there is an efficient load path directly from the inner casing structure 20 to the engine core 16 and through the pylon to the aircraft. Also, when the engine takes off and the engine is subject to intake upload, the bending loads imparted on the engine-are shared between the engine core and the C duct nacelle. The bifurcation regions of the C duct nacelle in particular are relatively rigid and resistant to such bending and therefore impart additional rigidity need to the engine core.

Because the discrete fastenings 36 are provided only in the bifurcation regions, they do not obstruct the pylon which is attached to the engine core. The discrete fastenings 36 are also lighter than an inner V groove and allow for the efficient transfer of load in the areas where is it required.

There is thus provided a gas turbine engine which includes an efficient means of transferring load from the C duct nacelle to the engine core.

Various modifications may be made to the above described embodiment without departing from the scope of the invention. In particular, the precise form of the discrete fastenings 36 may be modified as may their exact locations. The outer casing structure 22 of the nacelle 19 may be connected to the fan case 18 by an outer V groove or

by some other arrangement.

Referring to Fig. 6, in a further modification a rear lower bifurcation stabilisation link 46 may be provided, ball jointed onto the bottom of a tail bearing housing spoke structure (not illustrated) and engaged at the rear of the lower bifurcation wall 30 by closure of the C ducts in a similar manner to that explained in relation to the discrete fastenings 36. This would increase the structural effectiveness of the lower bifurcation wall 30. The upper bifurcation wall is commonly restrained against rotation when transferring axial loads from the core to the outer Vgroove by the radial restraint of the rear nacelle to pylon hinge location. For the lower bifurcation walls the rear load paths to the pylon are much longer and more flexible hence the lower bifurcation rotates more freely and is less effective at transferring axial shear from the core to the nacelle, unless a link to the tail bearing housing structure is provided. The link would be inclined, with its outer edge forward in order to balance out the axial and radial thermal growths of the core engine and the tail bearing housing structure. Alternatively, sliding contact surfaces could be utilised.

Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon.

Claims

CLAIMS 1. A ducted fan gas turbine engine (10) including: an engine core (16); a fan case (18) surrounding the core (16) and defining a bypass duct between the fan case (18) and the core (16); a C-duct assembly (19) downstream of the fan case (18) and including inner (20) and outer (22) generally annular casing structures defining a duct for bypass air therebetween; characterised in that the inner casing structure (20) of the C-duct assembly (19) is releasably attachable to the engine core via discrete fastenings (36).
2. A gas turbine engine according to Claim 2 wherein the outer casing structure (22) of the C-duct assembly (19) is attachable to the fan case via an annular projection, which engages an annular groove in the fan case.
3. A gas turbine engine according to Claim 1 or Claim 2 wherein the inner (20) and outer (22) casing structures of the C-duct assembly (19) together form a nacelle (19) which is split into two semi-annular halves, each half being hingably mountable on a pylon (31) of an aircraft, so that the two halves may hinge towards and away from one another.
4. A gas turbine engine according to Claim 3 wherein the discrete fastenings (36) are provided generally in the circumferential region of the nacelle which is hingably mountable on the pylon (31).
5. A gas turbine engine according to Claim 4 wherein further discrete fastenings (36) are provided in a region circumferentially opposite to this region.
6. A gas turbine engine according to any of Claims 3 to 5 wherein the discrete fastenings (36) are provided in a forward, upstream part of the nacelle (19).
7. A gas turbine engine according to any of Claims 3 to 6 wherein the nacelle (19) includes a bifurcation (28) joining the inner and outer casing structures generally in

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the region of the hinge attachment and a further bifurcation (30) in the region opposite thereto, and the discrete fastenings (36) are provided in the regions of the bifurcations (28,30).
8. A gas turbine engine according to Claim 7 wherein two discrete fastenings are provided in the region of each bifurcation.
9. A gas turbine engine according to any of Claims 3 to 8 wherein each discrete fastening (36) includes a pin (38) provided on the engine core (16) or on a member attached to the engine core (16) and a complementary recess (40) provided in the inner casing structure of the nacelle.
10. A gas turbine engine according to any of Claims 3 to 8 wherein each fastening (36) includes a pin (38) provided on the inner casing structure of the nacelle (19) and a complementary recess (40) provided on or associated with the engine core (16)..
11. A gas turbine engine according to Claim 9 or Claim 10 wherein the pins (38) are oriented with their axes tangential to the inner and outer casing structures (20, 22) and therefore parallel to the relative hinging movement between the inner casing structure and the engine core (16).
12. A gas turbine engine according to any of Claims 9 to 11 wherein each pin (38) includes a chamfered portion and each recess includes a complementary sloping portion. A single elongate member may constitute the pins for each of two fastenings
13. A gas turbine engine substantially as herein described with reference to the accompanying drawings.
14. Any novel subject matter or combination including novel subject matter disclosed herein, whether or not within the scope of or relating to the same invention as any of the preceding claims.