GB2588825A

GB2588825A - Method and assembly for securing payloads

Info

Publication number: GB2588825A
Application number: GB1916350.0A
Authority: GB
Inventors: Goulds Roberts
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2019-11-11
Filing date: 2019-11-11
Publication date: 2021-05-12
Also published as: GB201916350D0

Abstract

A method of securing a plurality of payloads 101 to a bracket 105 is disclosed, comprising an externally threaded elongate support member 107, two spring clips and an elongate spacer 112. The first spring clip 106 is mounted in an open state with a payload to the elongate support member. The clip is closed to secure the payload and an elongate spacer (108, fig6) is mounted to the elongate support member. The spacer is threaded on at least a portion of its interior surface to interface with the thread on the elongate support member and retain the spring clip in the closed state. A second spring clip 109 is then mounted to the elongate member with a second payload. The assembly for securing the payloads to the bracket is also disclosed. Additional spacers 112 and spring clips 111 could be used.

Description

METHOD AND ASSEMBLY FOR SECURING PAYLOADS

Field of the disclosure

The present disclosure relates to a method and assembly for securing payloads, in particular a method and assembly for securing payloads to a gas turbine engine, as well as a gas turbine engine including such an assembly.

Background of the Disclosure

Gas turbine engines typically have a number of payloads attached to various parts of the engine, such as the fan case and the core of the engine. For example, pipes, electrical harness and other payloads may be mounted to the engine using a bracket. Where several payloads are in close proximity to each other, they may be attached to a single bracket using spring clips, with spacers to separate them.

The spring clips are typically biased so that they are by default in an open position and can be held shut to be in a closed position so as to secure a payload. Where several payloads are to be secured to the engine in close proximity to each other, a "clip stack" can be used. A clip stack is typically constructed by joining multiple spring clips together to join them to the bracket and thus to secure the plurality of payloads.

In a clip stack, a plurality of spacers may be used to maintain the spring clips at the correct distance from each other. During assembly, the spacers must be held in position to prevent the spring clips from opening. This may cause difficulty during the assembly process.

It is an aim of the present disclosure to provide an improved method and assembly for securing payloads.

Summary of the Disclosure

According to a first aspect there is provided a method of securing a plurality of payloads to a bracket comprising providing on the bracket an elongate support member having a thread on its exterior surface, mounting a first spring clip in an open state to the elongate support member, mounting a first payload in the first spring clip, putting the first spring clip in a closed state to thereby secure the first payload, mounting a first elongate spacer on the elongate support member, the first elongate spacer being threaded on at least a portion of its interior surface such that the thread on the first elongate spacer interfaces with the thread on the elongate support member to thereby retain the first spring clip in the closed state, mounting a second spring clip in an open state to the elongate support member, mounting a second payload in the second spring clip, putting the second spring clip in the closed state to thereby secure the second payload.

The method may further comprise mounting a second elongate spacer on the elongate support member, the second elongate spacer being threaded on at least a portion of its interior surface such that the thread on the second elongate spacer interfaces with the thread on the elongate support member to thereby retain the second spring clip in the closed state.

The method may further comprise mounting at least one further spring clip in an open state on the elongate support member, mounting a further payload in each further spring clip, and putting each further spring clip in the closed state to thereby secure each further payload.

The method may further comprise mounting at least one further elongate spacer on the elongate support member, wherein each elongate spacer is threaded on at least a portion of its interior surface such that the thread on each elongate spacer interfaces with the thread on the elongate support member to thereby retain a respective spring clip in the closed state.

At least one elongate spacer may be threaded along the entirety of its interior surface.

At least one elongate spacer is threaded along a partial portion of its interior surface.

The elongate spacer may be mounted on the elongate support member such that the partially threaded portion is at the opposite end of the elongate spacer to the respective spring clip which it secures in the closed state.

The interior of at least one elongate spacer may have a substantially cylindrical shape.

The ratio of the axial length to the diameter of the substantially cylindrical shape may be at least 2:11 optionally at least 3:1 and optionally at least 4:1.

The exterior surface of at least one elongate spacer may be arranged to be engaged with a tool used to secure the spacer to the clip stack.

The exterior surface of at least one elongate spacer may have a hexagonal cross-sectional shape.

The elongate support member may extend through the bracket, and the method may further comprise fixing the position of the bracket relative to the elongate support member by mounting an elongate spacer on a first side of the bracket and mounting a fixing member a second side of the bracket.

The fixing member may be an elongate spacer.

In another aspect, there is provided an assembly for securing a plurality of payloads to a bracket comprising an elongate support member having a thread on its exterior surface, first and second spring clips mounted on the elongate support member and configured to be manipulated between an open state and a closed state, wherein the spring clips are configured to secure a payload when in the closed state and to allow insertion and removal of the payload when in the open state, and a first elongate spacer mounted on the elongate support member between the first and second spring clips, wherein the first elongate spacer is threaded on at least a portion of its interior surface such that the thread on the elongate spacer interfaces with the thread on the elongate support member to thereby retain the first spring clip in the closed state.

The assembly may further comprise a second elongate spacer mounted on the elongate support member, the second elongate spacer being threaded on at least a portion of its interior surface such that the thread on the second elongate spacer interfaces with the thread on the elongate support member to thereby retain the second spring clip in the closed state.

The assembly may further comprise at least one further spring clip mounted on the elongate support member and configured to be manipulated between an open state and a closed state, wherein the clips is configured to secure a payload when in the closed state and to allow insertion and removal of the payload when in the open state.

The assembly may further comprise at least one further elongate spacer mounted on the elongate support member between two of the spring clips, wherein each further elongate spacer is threaded on at least a portion of its interior surface such that the thread on each elongate spacer interfaces with the thread on the elongate support member to thereby retain a respective spring clip in the closed state.

At least one elongate spacer may be threaded along a partial portion of its interior surface.

The elongate spacer may be mounted such that the partially threaded portion is at the opposite end of the elongate spacer to the respective spring clip which it secures in the closed state.

The elongate support member may extend through the bracket, and an elongate spacer may be provided on a first side of the bracket, and a fixing member may be provided on a second side of the bracket, to thereby fix the position of the bracket relative to the elongate support member.

The fixing member may be an elongate spacer.

In another aspect, there is provided a gas turbine engine for an aircraft comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, a fan located upstream of the engine core, the fan comprising a plurality of fan blades, and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein the gas turbine engine comprises an assembly as defined above.

In the above gas turbine engine, the turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft, the engine core further may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine may comprise a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axially 25 downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used.

For example, the gearbox may be a "planetary" or "star" gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, 01 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a "star" gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform. The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity U. The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Ut1p2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and lhp is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph being Jkg-1K-1/(ms-1)2) The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the engine core. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg-ls, 105 Nkg-ls, 100 Nkg-ls, 95 Nkg-ls, 90 Nkg-ls, 85 Nkg-ls or 80 Nkes. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 80 Nkg-ls to 100 Nkg-ls, or 85 Nkg-ls to 95 Nkg-ls. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330kN to 420 kN, for example 350kN to 400kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3kPa, temperature 30 degrees C), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 1800K to 1950K. The maximum TEl may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.

As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this regard, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint -in terms of time and/or distance-between top of climb and start of descent. Cruise conditions thus define an operating point of, the gas turbine engine that provides a thrust that would ensure steady state operation (i.e. maintaining a constant altitude and constant Mach Number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example, where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide -in combination with any other engines on the aircraft -steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000 m to 15000 m, for example in the range of from 10000 m to 12000 m, for example in the range of from 10400 m to 11600 m (around 38000 ft), for example in the range of from 10500 m to 11500 m, for example in the range of from 10600 m to 11400 m, for example in the range of from 10700 m (around 35000 ft) to 11300 m, for example in the range of from 10800 m to 11200 m, for example in the range of from 10900 m to 11100 m, for example on the order of 11000 m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30kN to 35kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m). Purely by way of further example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50kN to 65kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000 ft (10668 m) In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.

According to an aspect, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gas turbine engine as described and/or claimed herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein. The operation according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein.

The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

Brief description of the drawings

Embodiments will now be described by way of example only, with reference to the Figures, in which: Figure 1 is a sectional side view of a gas turbine engine; Figure 2 is a close up sectional side view of an upstream portion of a gas turbine engine; Figure 3 is a partially cut-away view of a gearbox for a gas turbine engine; Figure 4 shows a bracket and a plurality of payloads; Figure 5 shows a bracket, a plurality of payloads, a first spring clip and a first elongate member; Figure 6 shows the arrangement of Figure 5, with the addition of an elongate spacer, Figure 7 shows the arrangement of Figure 6, with the addition of a second spring clip; Figure 8 shows the arrangement of Figure 7, with the addition of a second spacer; Figure 9 shows the arrangement of Figure 8, with a further spring clip; Figure 10 shows the arrangement of Figure 9, with a further elongate spacer; Figure 11 shows the arrangement of Figure 10 with a yet further spring clip; Figure 12 shows the arrangement of Figure 11 with a nut securing the final spring clip; Figure 13 shows an arrangement with an elongate support member extending through the bracket, and an elongate spacer on one side of the bracket and a fixing member on the opposite side of the bracket; Figure 14 shows various arrangements of partially threaded elongate spacers; and Figure 15 shows various arrangements of fully threaded spacers.

Detailed Description of the Disclosure

Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art.

Figure 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the core exhaust nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Figure 2. The low pressure turbine 19 (see Figure 1) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

Note that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as the "intermediate pressure turbine" and "intermediate pressure compressor". Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail in Figure 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in Figure 3. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in Figures 2 and 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the Figure 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of Figure 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in Figure 2.

Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in Figure 1 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core exhaust nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in Figure 1), and a circumferential direction (perpendicular to the page in the Figure 1 view). The axial, radial and circumferential directions are mutually perpendicular.

According to the present disclosure, there is provided a method of securing a plurality of payloads 101, 102, 103, 104 to a bracket 105. Such an arrangement may be used, for example, in a gas turbine engine, where payloads, such as pipes and/or electrical harness bundles, are secured to a part of the engine.

Typically, in such an arrangement, the bracket 105 is secured to the engine, and the various payloads are then secured to the bracket by constructing a clip stack. A clip stack is an assembly comprising a plurality of clips which can secure multiple payloads to a bracket.

As shown in Figure 4, there may be four payloads 101, 102, 103, 104, the payloads being shown in cross section in Figure 4. However, it will be appreciated that any combination or arrangement of payloads may be used. Likewise, the bracket 105 shown in figure 4 has a substantially rectangular cross section with a through-hole. However, it will be understood that the brackets may be of any suitable shape and form.

In a first stage of assembly, an elongate support member 107, such as a bolt, may be mounted to the bracket by passing it through a through hole in the bracket. The elongate support member 107 has a thread on its exterior surface. Alternatively, the elongate support member 107 may be integrally formed with the bracket.

Then, a first spring clip 106 is mounted to the elongate support member. As shown in Figure 5, the first spring clip 106 may have a main section for securing the payload, and two ends which can be attached to the elongate support member (for example by means of through-holes in the ends of the spring clip 106). Thus, the first spring clip is mounted to the elongate support member 107 by attaching its ends to the elongate support member 107. Simultaneously, or after this step, the first payload 101 is mounted in the first spring clip. For example, the first payload 101 may be mounted in the first spring clip by first attaching one end of the spring clip to the elongate support member, then inserting the payload into the clip, then attaching the second end of the spring clip to the elongate support member.

In Figure 5, the first spring clip 106 is in an open state, in which the payload can be inserted into, or removed from, the spring clip 106. The first spring clip can be manipulated between an open state and a closed state, with the payload 101 being securely held by the first spring clip 106 when in the closed state. The spring clip may be arranged such that it is made from a sprung material, which is biased towards being in the open position. Then, the spring clip can be manipulated into the closed position by applying a force to the spring clip. For example, if the two ends of the spring clip shown in Figure 5 are held together, the spring clip moves to the closed position.

In the next stage of assembly, the first spring clip 106 is moved to the closed position, and a first elongate spacer 108 is mounted on the elongate support member 107, as shown in Figure 6. The elongate spacer allows the distance between the first spring clip 106 and subsequent components in the clip stack (such as further spring clips) to be set. That is, the distance can be set by choosing a spacer with an appropriate length.

The first elongate spacer 108 has a thread on at least part of its inner surface. Thus, the first elongate spacer 108 can be mounted on the elongate support member 107 by screwing or threading the first elongate spacer 108 on to the elongate support member 107. When the end of the first elongate spacer reaches the ends of the first spring clip, it can apply a force to retain the first spring clip in the closed state. In other words, the first elongate spacer holds the first spring clip closed.

Second and subsequent payloads can then be added to the clip stack by adding further spring clips, and, if appropriate, further elongate spacers.

For example, in one arrangement, after the assembly in Figure 6 has been produced, a second spring clip 109 may be added, as in Figure 7, and a second payload mounted in the spring clip 109. Then, if only two payloads are present, and both payloads have been secured, the clip stack may be finished by adding a nut 110 (as shown in Figure 8) or other fastener (including, for example, another elongate spacer) to the end of the elongate support member 107.

It will be appreciated that, in some arrangements, further spring clips may be added in order to secure further payloads, and, in some arrangements, further elongate spacers may be used to secure each subsequent spring clip which is added to the clip stack in position. Any number of spring clips and spacers may be used according to the number of payloads which are to be secured to the clip stack. It will also be appreciated that two spring clips may be mounted to the elongate member without a spacer between them, so that they are at substantially the same axial position along the elongate support member. In such an arrangement, the two spring clips may be disposed at different circumferential positions around the axis of the elongate support member. Similarly, a conventional nut, rather than an elongate spacer, may be used between two spring clips, thus acting as a spacer.

The length of each spacer may be chosen so as to set the distance between each spring clip, and thus set the distance (or spacing) between each payload.

Figures 9-12 show further steps in the assembly of a clip stack using four spring clips to secure four payloads.

Figure 9 shows the addition of a third spring clip 111 to the arrangement of Figure 8. Namely, the third spring clip is mounted to the elongate support member 107 and the third payload 103 is mounted in the third spring clip 111. In this arrangement, the third spring clip 111 is mounted further along the elongate support member than the first and second spring clips, such that it is further from the bracket than the first and second spring clips.

Figure 10 shows the addition of a third elongate spacer 112 to the arrangement of Figure 9, with the third spring clip 111 in the closed position. Namely, the third elongate spacer 112 is mounted on the elongate support member 107. The third elongate spacer 112 holds the third spring clip 111 closed.

Figure 11 shows the addition of a fourth spring clip 113 to the arrangement of Figure 10. Namely, the fourth spring clip 113 is mounted to the elongate support member 107 and the fourth payload 104 is mounted in the fourth spring clip 113. The fourth spring clip 113 is mounted further along the elongate support member than the first, second and third spring clips, such that it is further from the bracket than the first, second and third spring clips.

Figure 12 shows the completion of the clip stack by mounting a nut 114 on the elongate support member 107 to secure the fourth spring clip 113.

It will be understood that, in the arrangement described above, the third and fourth spring clips, the third elongate spacer and the third and fourth payloads are examples of "further" spring clips, elongate spacers and payloads respectively.

In one arrangement, such as that shown in Figure 13, the elongate support member 107 may extend through the bracket 105 such that there is a part of the elongate support member extending from each end of the bracket. In such an arrangement, at least one elongate spacer 115 and another fixing member 116 may be used either side of the bracket in order to fix the position of the bracket relative to the elongate support member. The fixing member may be another elongate spacer, or may be a different type of fixing member, such as a conventional nut.

In this arrangement, a clip stack may be built up on either side of the bracket, and the desired number of clips and spacers can be mounted on each side of the bracket. Further, in such an arrangement, each side of the assembly (i.e. either side of the bracket) can be assembled and disassembled without requiring assembly or disassembly of the part on the other side of the bracket.

According to the present disclosure, various arrangements of elongate spacer may be used. Further, within a single clip stack, the elongate spacers need not be the same, but may be of different configurations (as described below) and of different lengths.

In some arrangements, and as shown in Figure 14, the elongate spacer may be threaded along only part of its interior surface 117 with the remainder 118 of the interior surface being unthreaded. Alternatively, as shown in Figure 15, an elongate spacer may be used which is threaded along the entirety of its interior surface.

The interior of the elongate spacers may have a substantially cylindrical shape. The ratio of the axial length of the cylindrical shape may be at least 2:1, at least 3:1, at least 4:1, or of any other suitable ratio.

The exterior surface of the elongate spacer may take various shapes. It may have a surface which allows it to be engaged with a tool used to secure the spacer to the clip stack. For example, the external surface may have a hexagonal cross-section, similar to a conventional nut. Alternatively, it may have a substantially cylindrical exterior surface, and include one or more flat portions, which may allow it to be manoeuvred.

When a partially threaded elongate spacer is used, the threaded portion may be towards one end of the elongate spacer. During assembly, the number of turns necessary to retain the clip in the closed position may be minimised by threading the elongate spacer onto the support member such that the threaded portion is at the opposite end to the clip which is being secured. Thus, it is not necessary to thread the elongate spacer along the entire distance between the end of the support member and the ends of the clip, but rather only along the distance until the other end of the spacer reaches the ends of the clip.

It will be understood that the invention is not limited to the embodiments above- described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

CLAIMS1. A method of securing a plurality of payloads (101, 102, 103, 104) to a bracket comprising: providing on the bracket (105) an elongate support member (107) having a thread on its exterior surface; mounting a first spring clip (106) in an open state to the elongate support member; mounting a first payload (101) in the first spring clip; putting the first spring clip in a closed state to thereby secure the first payload; mounting a first elongate spacer (108) on the elongate support member, the first elongate spacer being threaded on at least a portion of its interior surface such that the thread on the first elongate spacer interfaces with the thread on the elongate support member to thereby retain the first spring clip in the closed state; mounting a second spring clip (109) in an open state to the elongate support member; mounting a second payload (102) in the second spring clip; and putting the second spring clip in the closed state to thereby secure the second payload.
2 The method according to claim 1, further comprising mounting a second elongate spacer on the elongate support member, the second elongate spacer being threaded on at least a portion of its interior surface such that the thread on the second elongate spacer interfaces with the thread on the elongate support member to thereby retain the second spring clip in the closed state.
3 The method according to claim 1 or 2, further comprising: mounting at least one further spring clip (111, 113) in an open state on the elongate support member; mounting a further payload (103, 104) in each further spring clip; and putting each further spring clip in the closed state to thereby secure each further payload.
4. The method according to claim 3, further comprising: mounting at least one further elongate spacer (112) on the elongate support member; wherein each elongate spacer is threaded on at least a portion of its interior surface such that the thread on each elongate spacer interfaces with the thread on the elongate support member to thereby retain a respective spring clip in the closed state.
The method according to any one of the preceding claims, wherein at least one elongate spacer is threaded along the entirety of its interior surface.
6. The method according to any one of the preceding claims, wherein at least one elongate spacer is threaded along a partial portion of its interior surface.
7 The method according to claim 6, wherein the elongate spacer is mounted on the elongate support member such that the partially threaded portion is at the opposite end of the elongate spacer to the respective spring clip which it secures in the closed state.
8. The method according to any one of the preceding claims, wherein the interior of at least one elongate spacer has a substantially cylindrical shape.
9. The method according to claim 8, wherein the ratio of the axial length to the diameter of the substantially cylindrical shape is at least 2:1, optionally at least 3:1 and optionally at least 4:1.
10. The method according to any one of the preceding claims, wherein the exterior surface of at least one elongate spacer is arranged to be engaged with a tool used to secure the spacer to the clip stack.
11. The method according to any one of the preceding claims, wherein the exterior surface of at least one elongate spacer has a hexagonal cross-sectional shape.
12. The method according to any one of the preceding claims, wherein the elongate support member extends through the bracket, the method further comprising fixing the position of the bracket relative to the elongate support member by mounting an elongate spacer on a first side of the bracket and mounting a fixing member a second side of the bracket.
13. The method according to claim 12, wherein the fixing member is an elongate spacer.
14. An assembly for securing a plurality of payloads (101, 102, 103, 104) to a bracket (105) comprising: an elongate support member (107) having a thread on its exterior surface; first and second spring clips (106, 109) mounted on the elongate support member and configured to be manipulated between an open state and a closed state, wherein the spring clips are configured to secure a payload when in the closed state and to allow insertion and removal of the payload when in the open state; and a first elongate spacer (108) mounted on the elongate support member between the first and second spring clips; wherein the first elongate spacer is threaded on at least a portion of its interior surface such that the thread on the elongate spacer interfaces with the thread on the elongate support member to thereby retain the first spring clip in the closed state.
15. The assembly according to claim 14, further comprising a second elongate spacer mounted on the elongate support member, the second elongate spacer being threaded on at least a portion of its interior surface such that the thread on the second elongate spacer interfaces with the thread on the elongate support member to thereby retain the second spring clip in the closed state.
16. The assembly according to claim 14 or 15, further comprising at least one further spring clip (111, 113) mounted on the elongate support member and configured to be manipulated between an open state and a closed state, wherein the clips is configured to secure a payload when in the closed state and to allow insertion and removal of the payload when in the open state.
17. The assembly according to claim 16, further comprising at least one further elongate spacer (112) mounted on the elongate support member between two of the spring clips; wherein each further elongate spacer is threaded on at least a portion of its interior surface such that the thread on each elongate spacer interfaces with the thread on the elongate support member to thereby retain a respective spring clip in the closed state.
18. The assembly according to any one of claims 14-17, wherein at least one elongate spacer is threaded along the entirety of its interior surface.
19. The assembly according to any one of claims 14-18, wherein at least one elongate spacer is threaded along a partial portion of its interior surface.
20. The assembly according to claim 19, wherein the elongate spacer is mounted such that the partially threaded portion is at the opposite end of the elongate spacer to the respective spring clip which it secures in the closed state.
21. The assembly according to any one of claims 14-20, wherein the interior of at least one elongate spacer has a substantially cylindrical shape.
22. The assembly according to claim 21, wherein the ratio of the axial length to the diameter of the substantially cylindrical shape is at least 2:1, optionally at least 3:1 and optionally at least 4:1.
23. The assembly according to any one of claims 14-22, wherein the exterior surface of at least one elongate spacer is arranged to be engaged with a tool used to secure the spacer to the clip stack.
24. The assembly according to any one of claims 14-23, wherein the exterior surface of at least one elongate spacer has a hexagonal cross-sectional shape.
25. The assembly according to any one of claims 14-24, wherein the elongate support member extends through the bracket, and an elongate spacer is provided on a first side of the bracket, and a fixing member is provided on a second side of the bracket, to thereby fix the position of the bracket relative to the elongate support member.
26. The assembly according to claim 25, wherein the fixing member is an elongate spacer.
27.A gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft; wherein the gas turbine engine comprises an assembly according to any one of claims 13-26.