GB2045357A

GB2045357A - Aircraft turbine engine mounting assembly

Info

Publication number: GB2045357A
Application number: GB8006002A
Authority: GB
Original assignee: Westland Aircraft Ltd
Current assignee: Westland Group PLC
Priority date: 1979-03-06
Filing date: 1980-02-22
Publication date: 1980-10-29
Also published as: GB2045357B

Abstract

An aircraft turbine engine mounting assembly includes a plurality of mounts at least one of which 17, 25, 26 is arranged to provide a primary load path to support the engine when all the mounts are operative and a secondary load path to support the engine should one of the remaining mounts become inoperative. The mount for example comprises a lever 20 pivotally mounted at 21 intermediate its ends to airframe structure 37 and pivotally attached at 32 adjacent one end to the engine, the other end having an aperture 23 located around a pin 24 carried by the airframe structure, said pin having an outside diameter less than an inside diameter of the aperture, a resilient bush being located in an annular space between the surfaces of the pin and the aperture. No load is carried by the pin in the primary load path. <IMAGE>

Description

SPECIFICATION Aircraft turbine engine mounting assembly This invention relates to an aircraft turbine engine mounting assembly particularly but not exclusively for mounting a turbine engine on a helicopter.

Conventionally, turbine engines are attached to an aircraft by a mounting assembly including front and rear mounts extending between the engine and airframe structure. With such an arrangement, failure of either of the mounts may result in loss of the engine and maybe also the aircraft A prior arrangement for overcoming this problem is disclosed in United States Patent Specification Serial No: 3,907,220. In that arrangement, an additional redundant mount is provided at the rear mounting of a turbine engine, the redundant mount providing no support for the engine until one of the main mounts becomes inoperative. The requirement of an additional mount complicates the fitting and removal of the engine and results in an increase in the weight of the installation.

Accordingly, the present invention provides an aircraft turbine engine mounting assembly including a plurality of mounts connecting the engine to an airframe structure wherein at least one of the mounts is arranged to provide a primary load path to support the engine when all of the mounts are operative, and a secondary load path to support the engine should one of the remainder of the mounts become inoperative.

The mount may comprise a lever pivotally mounted intermediate its ends to the airframe structure for pivotal movement in a generally vertical plane, one end of the lever being pivotally attached to a mounting on the engine and the other end being associated with means arranged to limit pivotal movement of the lever to a predetermined amount Preferably, the means includes an aperture through the lever adjacent the end thereof and located around a pin carried by the airframe structure, the pin having an outside diameter less than an internal diameter of the aperture so that an annular space exists between the surfaces of the aperture and the pin when the pin is located centrally of the aperture.

In one form of the invention, a resilient bush is located in the annular space.

The intermediate pivotal attachment and the pivotal attachment to the engine may comprise spherical joints so as to permit pivotal movement of the lever in a generally horizontal plane about the intermediate attachment. Conveniently, in such an arrangement the apertured end of the lever may be located between spaced generally parallel flanges extending from the airframe structure, said flanges being spaced from side surfaces of the lever so as to limit the horizontal pivotal movement of the lever to a predetermined amount.

The end of the lever may be attached to the engine by oppositely directed tangential links each having one end pivotally attached to the lever and the other end pivotally attached to a peripheral flange on the engine.

The lever may be a straight lever and the centres of the two pivotal attachments and the aperture may be located on a common centreline.

Alternatively, the lever may be L-shaped with the intermediate pivotal attachment located at the junction of the two arms of the L-shaped lever.

In another aspect the invention provides an aircraft turbine engine mounting assembly including a front mounting and a rear mounting, wherein each mounting includes a mount comprising a lever pivotally mounted inteimediate its ends to airframe structure so as to be capable of limited pivotal movement in a generally vertical plane, the lever having an outer end pivotally attached to the engine and an apertured inner end, the aperture being located around a pin carried by the airframe structure, the pin having an outside diameter less than an inside diameter of the aperture whereby abutment of the surfaces of the pin and the aperture serves to limit pivotal movement of the lever to a predetermined amount.

In yet another aspect the invention provides an aircraft turbine engine mounting assembly including front and rear mountings. the front mounting comprising a spherical portion attached to one side of the engine on a horizontal centreline thereof and located in a mating socket attached to the airframe structure, and a horizontal lever pivotally mounted intermediate its ends to the airframe structure, an outer end of the lever being pivotally attached to a mounting on a lower surface of the engine and on a vertical centreline thereof, an inner end of the lever having an aperture located over a pin supported by the airframe structure and having an outside diameter less than an inside diameter of the aperture, the rear mounting comprising two L-shaped levers pivotally mounted intermediate their ends to the airframe structure, an outer end of one of the levers being attached to the engine by a pivotal attachment located at the side of the engine and on the horizontal centreline thereof and an outer end of the other lever being attached to the engine by a pivotal attachment on a lower surface of the engine and on the vertical centreline thereof, inner ends of both L-shaped levers having an aperture located over pins supported by the airframe structure, the pins having an outside diameter less than an inside diameter of its respective aperture.

The invention will now be described by way of example only and with reference to the accompanying drawings, in which.

Figure 1 is a side elevation of a helicopter showing the relative location of a turbine engine, Figure 2 is a simplified schematic illustration of the mounting assembly constructed in accordance with one embodiment of the invention showing front and rear mountings located on section lines X-X and Y-Y respectively of Figure 1, Figures 3 to 6 are schematic views similar to Figure 2 and illustrating the secondary load paths operative to support the engine in the event of a failure of any one of the mounts, and Figure 7 is a fragmentary view on an enlarged scaie taken in the direction of arrow D on Figure 2.

Referring now to Figure 1, a helicopter 11 includes a turbine engine (not shown) located within a compartment 12, the engine being attached to airframe structure by front and rear mountings located on lines X-X and Y-Y respectively.

In Figure 2 an engine 13, shown in outline only, is supported by a mounting assembly consisting of a front mounting 14 and a rear mounting 1 5.

Figure 2 includes an illustration of the loads effective on engine 1 3 during operation as consisting of horizontal loads (H), vertical loads (V), axial loads (A), and torque loads (T).

Front mounting 14 includes mount 16 and mount 17. Mount 16 is located at one side of engine 13 and on its horizontal centreline, and consists of a spherical portion 1 8 attached to the engine 1 3 and located in a mating socket 1 9 secured to the airframe.

Mount 1 7 (see also Figure 7) comprises a lever 20 pivotally mounted intermediate its ends on a spherical bearing 21 located between spaced parallel flanges 37 attached to the airframe. One end of the lever 20 is pivotally mounted on a spherical bearing 22 attached to a lower surface of the engine 13 on its vertical centreline. The other end of lever 20 is provided with an aperture 23 located around a pin 24 extending between the flanges 37. The internal diameter of the aperture 23 is greater than the external diameter of the pin 24 by a predetermined amount so that an annular space exists between the internal surface of the aperture 23 and the external surface of the pin.In the illustrated embodiment, a resilient bush 38 (Figure 7) is located in the space, the arrangement permitting limited pivotal movement of the lever 20 about the bearing 21 by compression of the resilient bush.

Rear mounting 1 5 includes mount 25 and mount 26. Mount 25 is located at one side of engine 13 and on its horizontal centreline and mount 26 is located beneath the engine and on its vertical centreline.

Mounts 25 and 26 are similar in construction, and include L-shaped levers 27 and 28 respectively. Lever 27 is pivotally mounted on a spherical bearing 29 attached to the airframe structure so that one arm extends horizontally towards the engine 1 3 on the horizontal centreline thereof. An outer end of this arm of lever 27 is pivotally attached to two oppositely directed tangential links 30 which are pivotally connected to a peripheral mounting flange on the engine 13.

The other arm of lever 27 extends vertically downwardiy to an end provided with an aperture 31 located around a pin 32 supported from the airframe structure and in a manner similar to that previously described in relation to mount 17, so as to permit limited pivotal movement in a vertical plane of lever 27 about the spherical bearing 27.

Lever 28 is similarly constructed, and is mounted on a spherical bearing 33 so that one arm extends vertically upwardly and on a vertical centreline of the engine 13 to support tangential links 34 attached to a mounting flange on the engine 13, and the other arm extends horizontally to an end provided with an aperture 35 located around a pin 36 in a manner similar to that previously described.

In the illustrated embodiment, the clearance between the inside diameter of the apertures 23, 31 and 35 and the outside diameter of the respective pins 24, 32 and 36 is 1,0 m.m.

approximately.

Having now described the various mounts comprising the illustrated embodiment of a mounting assembly of this invention it will be convenient to set out the loads carried by each mount during normal operation when all mounts are operational, and it is to be understood that, in this condition, the mounts 17, 25 and 26 are located such that the pins 24, 32 and 36 are located centrally of the respective apertures 23,31 and 35 so that the resilient bush in the space is not compressed and no load is carried at this end of the respective levers 20, 27 and 28.

Thus, the primary load paths illustrated by the full line arrows in Figure 2, of the respective mounts are as follows:- 1. Mount 16 Mount 1 6 caters for vertical (V), horizontal (H) and axial (A) loads by reaction between the spherical portion 18 and the socket 19, as well as torque loads (T) as a horizontal couple in conjunction with lever 20.

2. Mount 17 Mount 1 7 caters for torque loads (T) as a horizontal couple in conjunction with mount 1 6.

3. Mount 25 Mount 25 caters for horizontal loads (H) by compression or tension in the horizontal arm of lever 27.

4. Mount 26 Mount 26 caters for vertical loads (V) by compression of the vertical arm of lever 28.

The effect of a failure of any one of the mounts 16, 1 7, 25 or 26 will now be described with reference to Figures 3 to 6 inclusive in each of which a theoretically failed mount has been deleted. The forces normally taken by each of the respective mounts have previously been described so that the following description refers to the load paths operative in the event of failure. These load paths are illustrated by the broken line arrows in Figures 3 to 6 inclusive.

1. Failure of Mount 16 (Figure 3) The vertical (V) loads normally taken by mount 16 are now taken by vertical bending of lever 20, this load being reacted by force Ry provided by a secondary load path resulting from pivotal movement of the lever 20 about the bearing 21 causing maximum compression of the resilient bush between the surfaces of the aperture 23 and the pin 24.

Axial loads (A) are taken by horizontal bending of lever 20, this load being reacted by force RA provided by a secondary load path resulting from pivotal movement of the lever 20 about the bearing 21 in a horizontal plane causing an inner end of the lever 20 adjacent the aperture 23 to abut an inner surface of the supporting flanges 37.

This movement is permitted by the resilient bush between the surfaces of the pin 24 and the aperture 23 (refer also to Figure 7).

Horizontal loads (H) are taken by tensile or compressive loads in lever 20.

Torque loads (T) are taken by auxilary torque capability of the rear mounting 1 5 through horizontal couple B and/or vertical couple C as illustrated in Figure 3. Horizontal couple B is reacted by force R81 and RB2 the former being provided by compressive load in the horizontal arm of L-shaped lever 27. Force RB2 is provided by a secondary load path resulting from pivotal movement of the L-shaped lever 28 about the bearing 33 causing maximum compression of the resilient bush between the surfaces of the aperture 35 and the pin 36.

Vertical couple C is reacted by force Rc provided by a secondary load path resulting from pivotal movement of the L-shaped lever 27 about bearing 29 causing maximum compression of the resilient bush between the surfaces of aperture 31 and pin 32.

2. Failure of Mount 17 (Figure 4) Mount 1 6 operates in its normal capacity to take all vertical (V), axial (A) and horizontal (H) loads.

Torque loads (T) are taken by the auxiliary torque capability of the rear mounting 1 5 as previously described with reference to Figure 3.

3. Failure of Mount 25 (Figure 5) Horizontal loads (H) are now taken by mount 26 and are reacted by force RH provided by a secondary load path resulting from pivotal movement of L-shaped lever 28 about bearing 33 causing maximum compression of the resilient bush between the surfaces of aperture 35 and pin 36.

4. Failure of Mount 26 (Figure 6) Veritcal loads (V) are now taken by mount 25 and are reacted by force RV provided by a secondary load path resulting from pivotal movement of L-shaped lever 27 about bearing 29 causing maximum compression of the resilient bush between the surfaces of aperture 31 and pin 32.

Thus in the described embodiment, the present invention provides a mounting assembly comprising mounts 1 6, 1 7, 25 and 26 which mounts combine to provide primary load paths to support the engine 13 against vertical, axial, horizontal and torque loads, and in which any one of mounts 17, 26 and 25 are constructed so as to also provide various secondary load paths capable of providing continued support for the engine 1 3 in the event of a failure of any one of mounts 16, 17, 25 and 26.

In the case of all three of the mounts 17, 25 and 26, a secondary load path is provided by pivotal movement of the respective levers 20, 27 and 28 in a vertical plane generally perpendicular to a longitudinal axis of the engine 13 so as to provide a reaction force resulting from maximum compression of the resilient bushes between an inner surface of apertures 23, 31 and 35 and an outside diameter of respective pins 24, 32 and 36.

Furthermore, mount 17 is arranged to provide a further secondary load path by reaction between the sides of the apertured end of lever 20 and adjacent inner surfaces of mounting brackets 37.

It should be understood that the use of a straight lever such as lever 20 or L-shaped levers such as 27 and 28 is dictated by such considerations as available installation space, and that a similar facility to that previously described would be provided by any combination of straight and/or L-shaped levers at any of the mounting positions.

Claims

1. An aircraft turbine engine mounting assembly including a plurality of mounts connecting the engine to an airframe structure, wherein at least one of the mounts is arranged to provide a primary load path to support the engine when all the mounts are operative and a secondary load path to support the engine should one of the remainder of the mounts become inoperative.

2. A mounting assembly as claimed in Claim 1, wherein said mount comprises a lever pivotally mounted intermediate its ends to the airframe structure for pivotal movement in a generally vertical plane, one end of the lever being pivotally attached to a mounting on the engine and the other end being associated with means to limit pivotal movement of said lever to a predetermined amount

3. A mounting assembly as claimed in Claim 2, wherein said limiting means includes an aperture through the lever adjacent the end thereof and located around a pin carried by the airframe structure, said pin having an outside diameter less than an internal diameter of the aperture so that an annular space exists between the surfaces of the pin and the aperture when the pin is located centrally thereof.

4. A mounting assembly as claimed in Claim 3, wherein a resilient bush is located in said annular space.

5. A mounting assembly as claimed in any preceding Claim, wherein said intermediate pivotal attachment and said pivotal attachment to the engine comprise spherical joints.

6. A mounting assembly as claimed in Claim 5, wherein the apertured end of the lever is located between generally parallel flanges attached to said airframe structure, said flanges being spacedapart from side surfaces of the lever.

7. A mounting assembly as claimed in any preceding Claim, wherein the lever is attached to the engine through oppositely directed tangential links each having one end pivotally attached to the lever and the other end pivotally attached to a peripheral flange on the engine.

8. Mounting assembly as claimed in any preceding Claim, wherein said lever is a substantially straight lever having the centres of the two pivotal attachments and the centre of the aperture located on a common centreline.

9. A mounting assembly as claimed in any one of Claims 1 to 9 inclusive, wherein said lever is an L-shaped lever having said intermediate pivotal attachment located at the junction of the two arms thereof.

1 0. An aircraft turbine engine mounting assembly including front and rear mountings, wherein each mounting includes a mount comprising a lever pivotally mounted intermediate its ends to airframe structure so as to be capable of limited pivotal movements in a generally vertical plane, said lever having one end pivotally attached to the engine and an apertured other end. the aperture being located around a pin carried by the airframe structure, said pin having an outside diameter less than an inside diameter of the aperture.

11. An aircraft turbine engine mounting assembly having front and rear mountings, the front mounting comprising a spherical portion attached to one side of the engine on a horizontal centreline thereof and located in a mating socket attached to the airframe structure, and a horizontal lever pivotally mounted intermediate its ends to the airframe structure, one end of the lever being pivotally attached to a mounting on a lower surface of the engine on a vertical centreline thereof and its other end having an aperture located over a pin supported by the airframe structure and having an outside diameter less than an inside diameter of the aperture, said rear mounting comprising two L-shaped levers pivotally mounted intermediate their ends to said airframe structure, an end of one of said levers being attached to the engine by a pivotal attachment located at the side of the engine and on a horizontal centreline thereof and an end of said other lever being attached to the engine by a pivotal attachment on a lower surface of the engine on a vertical centreline thereof, the other ends of both said L-shaped levers having an aperture located over a pin supported by the airframe structure, said pins having an outside diameter less than an inside diameter of its respective aperture.

12. An aircraft turbine engine mounting assembly substantially as herein described and illustrated in the accompanying drawings.

13. Every novel feature and every novel combination of features disclosed herein.