CN107429614A - Hybrid engine component with installation cage portion - Google Patents
Hybrid engine component with installation cage portion Download PDFInfo
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- CN107429614A CN107429614A CN201680023261.2A CN201680023261A CN107429614A CN 107429614 A CN107429614 A CN 107429614A CN 201680023261 A CN201680023261 A CN 201680023261A CN 107429614 A CN107429614 A CN 107429614A
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Classifications
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T10/00—Road transport of goods or passengers
- Y02T10/10—Internal combustion engine [ICE] based vehicles
- Y02T10/12—Improving ICE efficiencies
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- Structures Of Non-Positive Displacement Pumps (AREA)
- Supercharger (AREA)
Abstract
Hybrid engine component with engine core, including at least one explosive motor, turbine section and the compressor with the outlet being in fluid communication with the entrance of engine core.Shell is connected to turbine section, compressor and engine core.Installation cage portion is connected to the installation portion that shell is attached between compressor and thermal region, the thermal region include turbine section and（It is multiple）Blast pipe.Pillar is separated by least one fire wall with thermal region.Installation cage portion can include multiple pillars that turbine section and engine core are all extended outwardly away from from installation portion.Shell can be gearbox module shell, and turbine shaft is engaged by it with engine shaft.Installation cage portion can be completely contained in following axial space, wherein turbine section and（It is multiple）Blast pipe is located at the outside of the axial space.
Description
The cross reference of related application
This application claims in the U. S. application No. 15/047,362 submitted for 18 days 2 months in 2016 priority, it is in 2015
The part continuation application for the U. S. application No. 14/864,124 that on September is submitted for 24, it requires to submit on 2 20th, 2015
U.S. Provisional Application No. 62/118,914 priority, its whole entire content is incorporated herein by reference.
Technical field
The application relates generally to hybrid engine component, and relates more specifically to supercharging or the whirlpool used in aircraft
Wheel supercharging hybrid engine component.
Background technology
It is relatively large that volume can be limited including the hybrid engine component as booster or the compressor of turbocharger
Component, it is likely difficult to fit into existing aircraft nacelle, therefore generates for it is applied into aircraft applications
Some are difficult.
The content of the invention
In one aspect, there is provided hybrid engine component, including：Engine core, including engaged with engine shaft driving
At least one explosive motor；Turbine section, there is the outlet fluid by least one blast pipe and engine core
The entrance of connection, the turbine section include being connected at least one turbine rotor of turbine shaft, the turbine shaft
Engaged with engine shaft driving；Compressor, has the outlet being in fluid communication with the entrance of engine core, and the compressor includes
At least one compressor drum engaged with least one driving in turbine shaft and engine shaft；It is connected to turbine portion
The shell of section, compressor and engine core；And installation cage portion, it is described for hybrid engine component to be attached into aircraft
Installation cage portion includes being connected to multiple pillars of installation portion, and the installation portion is attached to shell, the hot-zone on the outside of thermal region
Domain includes the adjacent with least one blast pipe of turbine section, at least one blast pipe and engine core
Part, the pillar are separated by least one fire wall with thermal region.
On the other hand, there is provided hybrid engine component, including：Engine core, including engaged with engine shaft driving
At least one explosive motor；Gearbox module, including the gearbox module shell comprising at least one pinion unit；Turbine
Machine section, in the outside of gearbox module shell, the turbine section, which has, passes through at least one blast pipe and engine core
Communication entrance, the turbine section includes being connected at least one turbine rotor of turbine shaft, institute
One stated at least one pinion unit that turbine shaft passes through gearbox module engages with engine shaft driving；Pressure
Contracting machine, in the outside of gearbox module shell, the compressor has the outlet being in fluid communication with the entrance of engine core, described
Compressor includes at least one compressor drum engaged with least one driving in turbine shaft and engine shaft；Wherein,
Turbine section and engine core are located on the identical sidepiece of gearbox module shell, and compressor is located at gearbox module
On the relative sidepiece of shell；And installation cage portion, for hybrid engine component to be attached into aircraft and is connected to shell, institute
Installation cage portion is stated to be kept completely separate by least one fire wall and turbine section and at least one blast pipe.
Brief description of the drawings
With reference now to accompanying drawing, wherein：
Fig. 1 is the schematic diagram according to the hybrid engine component of specific embodiment；
Fig. 2 is the wankel that can be used in all hybrid engine components as shown in Figure 1 according to specific embodiment
（Wankel）The viewgraph of cross-section of engine；
Fig. 3 is the schematic three-dimensional views according to Fig. 1 of specific embodiment hybrid engine component；
Fig. 4 is the schematic side elevation according to Fig. 3 with engine mount of specific embodiment hybrid engine component；
Fig. 5 is that have entry conductor and fire wall according to specific embodiment（firewall）Fig. 3 hybrid engine component
Cross-sectional schematic side view；
Fig. 6 is the front schematic view according to Fig. 3 of specific embodiment hybrid engine component；
Fig. 7 is the schematic diagram according to the hybrid engine component of another specific embodiment；
Fig. 8 is the schematic three-dimensional views according to Fig. 7 of specific embodiment hybrid engine component；
Fig. 9 is the schematic horizontal stroke according to Fig. 8 with entry conductor and fire wall of specific embodiment hybrid engine component
Side cross-sectional view；
Figure 10 A are the schematic three-dimensional views according to Fig. 8 with engine mount of specific embodiment hybrid engine component；
Figure 10 B are Figure 10 A hybrid engine component and the schematic side elevation of engine mount；
Figure 11 is the schematic exploded end view according to the hybrid engine component of another specific embodiment；And
Figure 12 is the schematic side elevation of the part of Figure 11 hybrid engine component.
Embodiment
With reference to figure 1, hybrid engine component 10 is generally illustrated, includes the more rotor rotary engines of heavy fuel of liquid cooling
Core 12.Engine core 12 has engine shaft 16, and the engine shaft 16 is driven and driven by engine core 12 can
Rotary load, it is shown in which as drive shaft 8.Drive shaft 8 can be the integrally formed part of engine shaft 16, be directly connected to
To it or pass through gear-box（It is not shown）It is connected to it.It should be understood that hybrid engine component 10 can be alternately configured to
The load of any suitable type is driven, is included but is not limited to, one or more generators, propeller, annex, rotor mast
（rotor mast）, compressor or the load of any other suitable type or its combination.
Hybrid engine component 10 is configured to single-shaft engine.Term " single shaft " is intended to describe all rotating parts herein
Part（（It is multiple）Compressor drum,（It is multiple）Turbine rotor, engine shaft, annex）Directly or through one or more
The hybrid engine that gear-box mechanically links together.Therefore, " single shaft " engine may include two or more mechanical chains
The axle connect.Term " single shaft " is intended to and has two or more spools（spool）Engine be contrasted, it is described two or
More spools can rotate freely so as to include one or more freedom turbines relative to each other.
Hybrid engine component 10 includes the entrance that compressed air is fed to engine core 12（With engine core 12
Each engine ingress port is corresponding or connection）Compressor 14.Engine core 12 receives pressurization sky from compressor 14
Gas, and at elevated pressures burning fuel to provide energy.Mechanical output drives engine shaft as caused by engine core 12
16.Each engine of engine core 12 provides exhaust stream, and the form of the exhaust stream is the height left with high peak velocity
Press the exhaust pulses of hot gas.The outlet of engine core 12（With the exhaust port pair of each engine of engine core 12
Should or it connect）It is in fluid communication with the entrance of turbine section 18, and the exhaust stream therefore from engine core 12 is supplied
To turbine section 18.Turbine section 18 drive compressor 14 and with the hybrid power of engine shaft 16.
In a particular embodiment, hybrid engine component includes four（4）Primary module：Include the core of engine core 12
Module 22, the cold section of gearbox module 20 including compressor 14 or compressor module 24 and including turbine section 18
Hot section or turbine module 28.In a particular embodiment, turbine module 28 and compressor module 24 can be at the scene by typical cases
Attendant remove, while hybrid engine component 10 remains attached to aircraft, in order to maintenance, repair and/or replacing.
In a particular embodiment, turbine module 28, compressor module 24 and nucleus module 22 can in a manner of independent and separation from
Hybrid engine component 10 is dismantled and removed, i.e., need not dismantle/remove any other module；In a particular embodiment, each
Therefore the part of module is comprised in shell and/or is mounted to shell, the shell limits encapsulation independently of the part of other modules
Part.In a particular embodiment, the modularity of hybrid engine component 10 can allow to reduce or minimize hybrid engine component 10
In part number and/or may be such that each module can be run under the speed corresponding to optimal performance condition.
With reference to figure 3, nucleus module 22 includes engine core 12 and fuel dispensing system 13.In the illustrated embodiment,
Engine core 12 includes being drivingly engaged to multiple rotary engine 12' of axle 16, and fuel dispensing system 13 includes feedback
Send the guide for each rotary engine（pilot）With the public common-rail road 13' of main ejector.Although engine core 12
It is depicted as including two rotary engine 12', it should be understood that in another embodiment, engine core 12 may include more than two
Rotary engine 12'（Such as 3 or 4 rotary engines）, or single rotary engine 12'.Each rotary engine 12'
With the rotor being sealingly engaged in respective housings, wherein each rotary engine 12' has approximately constant volume combustion rank
Section, for high cycle efficieny.In the illustrated embodiment, each rotary engine 12' is Wankel engine.
With reference to figure 2, the example for the Wankel engine that can be used as rotary engine 12' in engine core 12 is shown
Property embodiment.Each Wankel engine 12' includes the housing 32 for limiting inner chamber, and the inner chamber, which has, limits two projections
Portion（lobe）Profile, it is preferably epitrochoid（epitrochoid）.Rotor 34 is received in inner chamber.Rotor limits
Fixed three circumferentially spaced tops（apex）Part 36, and the profile of the general triangular of the sidepiece with outside archwise.
The inner surface of top section 36 and the periphery wall 38 of housing 32 is sealingly engaged, to form three between rotor 34 and housing 32
Work chamber 40.
Rotor 34 is bonded to the eccentric part 42 of axle 16, so that internally intracavitary performs revolution（orbital
revolution）.Axle 16 performs three rotations of each revolution for rotor 34.The geometrical axis 44 of rotor 34 is from housing 32
Axis 46 offset and parallel with it.During each revolution, the Volume Changes of each chamber 40, and moved around inner chamber
It is dynamic, to undergo air inlet, compression, expansion and exhaust four-stage.
Air inlet port 48 is provided through periphery wall 38, for allowing compressed air to enter each work chamber 40 successively.
Exhaust port 50 is also provided through periphery wall 38, for successively discharging exhaust gas from each work chamber 40.For electricity
Heat plug（glow plug）, spark plug or the passage 52 of other igniting elements and fuel injector be also provided through periphery wall
38.Alternatively, air inlet port 48, exhaust port 50 and/or passage 52 can be provided through end or the side wall 54 of housing；
And/or igniting element and pilot fuel injection device can be with the first guide chambers that are limited in housing 32 and are connected with inner chamber
Room（It is not shown）Connection, for providing pilot injection.Guide sub-chamber can for example be limited at reception in periphery wall 38
Insert（It is not shown）In.
In the fig. 3 embodiment, fuel injector is common-rail fuel injector, and and heavy fuel（For example, bavin
Oil, kerosene（Aviation fuel）, equivalent biological fuel）Source connection, and heavy fuel is sent to（It is multiple）In engine so that
Combustion chamber is layered, wherein being fuel-rich-air mixture near incendiary source and being barrenr mixed in other places
Compound.
Referring again to Fig. 2, in order to effectively operate, work chamber 40 is sealed, for example, by from rotor 34 extend so as to
The spring-loaded top seal 56 of periphery wall 38 is engaged, and by extending from rotor 34 so as to the bullet of engagement end portion wall 54
Spring loading surface or seals 58 and end or corner sealing part 60.It is close that rotor 34 also includes at least one spring-loaded oil
Seal ring 62, it surrounds the bearing for the rotor 34 being used on eccentric shaft part 42 and biased against end wall 54.
Each Wankel engine provides the exhaust stream that form is relatively long exhaust pulses；For example, in specific embodiment
In, 360 ° rotations of each Wankel engine per secondary axes, which have, once explodes, wherein pact of the exhaust port for the rotation
270 ° stay open, therefore provide about 75% pulse duty factor.In contrast to this, the piston of reciprocating 4 reciprocating piston engine
Generally per 720 ° of secondary axes, rotation, which has, once explodes, and wherein exhaust port stays open for about 180 ° of the rotation, therefore
25% pulse duty factor is provided.
Especially but not exclusively it can be suitable in the specific embodiment of low altitude area, each Wankel engine has
From 5 to 9 volume expansion ratio, and follow Miller cycle（Miller cycle）Operation, have lower than volume expansion ratio
Volume compression ratio, such as by the way that air inlet port is positioned to compared with the volume compression engine equal or similar with expansion ratio more
Close to dieback point（TDC, top dead center）.Alternatively, each Wankel engine can be with similar or identical volume
Compress with expansion ratio to operate.
It should be understood that other configurations are possible for engine core 12.Engine core 12（It is multiple）Start
Machine 12' configuration, for example, the placement of port, the quantity of seal and placement, the quantity etc. of fuel injector, can be with showing
Embodiment it is different.In addition, it should be understood that each engine 12' of engine core 12 can be the internal combustion of any other type
Engine, the rotary engine of including but not limited to any other type, and the non-rotating internal-combustion engine of any other type
Machine, such as reciprocating engine.
Referring again to Fig. 1, in a particular embodiment, compressor 14 is the centrifugal compressor for having single impeller 14'.It is other
Configuration is alternatively possible.Compressor 14 can be single-stage system or multilevel device, and may include there is circumferentially array
Radially, axially or mixed flow movable vane（blade）One or more rotors.
With reference to figure 3, gearbox module 20 includes shell 21, and the shell 21 includes（For example, encapsulation）At least one gear drive
System, and compressor module 24 includes the shell 25 positioned at the outside of gearbox module shell 21.Compressor module shell 25 includes（Such as seal
Dress）（It is multiple）Compressor drum 14'（Such as impeller）, diffuser（diffuser）, cover（shroud）, inlet volute（inlet
scroll）And variable inlet stator（guide vane）88（See Fig. 1）, air reaches（It is multiple）Before compressor drum
Cycle through the variable inlet stator 88.Compressor module shell 25 can include coordinating to limit the encapsulation for including compressor 14
Multiple shells of part, and/or can be limited in whole or in part by the outer wall of compressor 14.With reference to figure 3-4, compressor module
Shell 25 is installed on the face of gearbox module shell 21.In a particular embodiment, compressor module shell 25 and gearbox module shell
21 are removably interconnected, such as are bolted and/or clamping part or by using any other by making the adjacent flanges of shell 25,21
The fastener of suitable type（Including but not limited to, limit and be known as " quickly accessing disconnection（quick access
disconnect）" connection type such engagement member or fastener）Interconnection.Other configurations are also possible.
With reference to figure 4, in a particular embodiment, the connection between the outlet of compressor 14 and the entrance of engine core 12 is led to
Inlet manifold 15 is crossed to perform.In a particular embodiment,（It is multiple）Compressor drum is sized to supply engine quality
Stream and main cabin bleed（cabin air bleed）.It may be configured as the inlet manifold 15 separated with compressor module 24, including branch
Port 15', the main cabin bleed for pressurization.
Turbine module 28 include comprising（Such as encapsulate）The turbine module shell 29 of turbine section 18, the turbine
Section 18 includes being connected at least one rotor of turbine shaft 19, has corresponding（It is multiple）Turbine vane（vane）、（It is more
It is individual）Housing,（It is multiple）Include feature（containment feature）And（It is multiple）Tie bolt.Turbine module shell 29
It is spaced apart with compressor module shell 25, and also is located at the outside of gearbox module shell 21.Turbine module shell 29 can include
Coordinate to limit the multiple shells for the packaging part for including turbine section 18, and/or the outer wall of turbine section 18 can be passed through
Limit in whole or in part.Turbine module shell 29 is arranged on the face with receiving compressor module shell 25 of gearbox module shell 21
On relative face；In a particular embodiment, turbine module shell 29 is installed on the face forward of gearbox module shell 21.
In specific embodiment, turbine module shell 29 and gearbox module shell 21 are removably interconnected, such as by making the neighbour of shell 29,21
Flange is connect to be bolted and/or clamping part or the fastener by using any other suitable type（Including but not limited to, limit
Know the such engagement member or fastener of the connection type for " quickly accessing disconnection "）Interconnection.Other configurations are also possible.
Multiple blast pipes 30 provide the outlet of engine core 12（Each engine 12' exhaust port）And turbine portion
Fluid communication between the entrance of section 18.Nucleus module 22 is installed in the identical with turbine module 28 of gearbox module shell 21
Face on；In a particular embodiment, turbine module 28 to the close-coupled of nucleus module 22 by by the He of engine core 12
Blast pipe 30 between turbine section 18 remains as short as possible and controls the flow region run through, so as to help to increase
（And preferably maximize）Exhaust gas energy regenerating.Blast pipe 310 becomes very hot during use, and therefore implements
Suitable material selects and cooling, to ensure its durability.
As that can see in Fig. 1, turbine section 18 may include one or more included in turbine module shell
Individual turbine stage.In a particular embodiment, turbine section 18 includes the first order for receiving the exhaust from engine core 12
Turbine 26, and receive the second stage turbine 27 of the exhaust from the first stage turbine 26.First stage turbine 26 by with
It is set to speed mode turbine（velocity turbine）, also known as tangential turbine（impulse turbine）, and
The kinetic energy of core exhaust gas is reclaimed, while is produced to the minimum back pressure of the exhaust of engine core 12 or no back pressure.
Second stage turbine 27 is configured to pressure type turbine（pressure turbine）, also known as reaction turbine
（reaction turbine）, and complete the recovery of the available mechanical energy from exhaust gas.Each turbine 26,27
Can be centrifugation or the axial device with one or more rotors, one or more rotor has circumferentially array
Radially, axially or mixed flow movable vane.In another embodiment, turbine section 18 may include single turbine, be configured to
Tangential turbine or pressure type turbine.
Pure tangential turbine is by changing the direction of stream without accelerating the stream on the inside of rotor come work；Fluid it is deflected and
The obvious pressure drop of rotor movable vane is not crossed over.The movable vane of pure tangential turbine is designed such that perpendicular to flowing side
To transverse plane in, edge of the area in movable vane limited between movable vane and be identical at the trailing edge in movable vane：Turbine
The flow area of machine is constant, and movable vane is symmetrical generally about the plane of rotating disk.Pure tangential turbine is done
Work(is only because the change in the direction of the stream by turbine bucket.Typical pure tangential turbine includes steam and hydraulic pressure whirlpool
Turbine.
In contrast, reaction turbine accelerate rotor on the inside of stream, but need across rotor static pressure drop so that
The stream must be can speed up.The movable vane of reaction turbine is designed such that in the transverse plane perpendicular to flow direction, moves
The area limited between leaf is bigger than at the trailing edge of movable vane in the edge of movable vane：The flow area streamwise of turbine
Reduce, and movable vane is not symmetrical generally about the plane of rotating disk.At least part of the work(done of pure reaction turbine is
Due to the acceleration of the stream by turbine bucket.
Most of aeroturbines are not " pure impulses " or " pure reactionary style ", but it is opposite but complementary to follow the two
Principle mixing and operate --- i.e., the pressure drop across movable vane, the flow area streamwise of turbine bucket be present
In the presence of some reductions, and the speed of the rotation of turbine is due to the direction change of stream and accelerates the two.The reaction of turbine
Degree（degree of reaction）The reaction ratio based on temperature can be used（Equation 1）Or the reaction ratio based on pressure（Equation
2）To determine, it is generally numerically closer to each other for identical turbine：
Wherein, T is temperature and P is pressure, and s refers to static side mouth, and numeral refers to the measurement position of temperature or pressure：0 is used for turbine
Machine stator blade（Stator）Entrance, 3 are used for turbine bucket（Rotor）Entrance, and 5 are used for turbine bucket（Rotor）Go out
Mouthful；And wherein pure tangential turbine can have 0（0%）Ratio, and pure reaction turbine can have 1（100%）'s
Ratio.
In a particular embodiment, the first stage turbine 26 is configured to utilize the stream of pulses for leaving engine core 12
（pulsating flow）Kinetic energy, while make stream stable, and the second stage turbine 27 is configured to from the residual compression in stream
Energy is extracted, while expands stream.Therefore, the first stage turbine 26 has more anti-than smaller than the reaction of the second stage turbine 27
Dynamic ratio.
In a particular embodiment, the second stage turbine 27 has the reaction ratio higher than 0.25；In another specific embodiment,
Second stage turbine 27 has the reaction ratio higher than 0.3；In another specific embodiment, the second stage turbine 27 has about 0.5
Reaction ratio；In another specific embodiment, the second stage turbine 27 has the reaction ratio higher than 0.5.
In a particular embodiment, the first stage turbine 26 has most 0.2 reaction ratio；In another specific embodiment,
First stage turbine 26 has most 0.15 reaction ratio；In another specific embodiment, the first stage turbine 26 has most
0.1 reaction ratio；In another specific embodiment, the first stage turbine 26 has most 0.05 reaction ratio.
It should be understood that any suitable reaction ratio for the second stage turbine 27（Including but not limited to, in above-mentioned reaction ratio
Any one）Can with for the first stage turbine 26（Including but not limited to, any one of above-mentioned reaction than in）Appoint
What suitable reaction is than combination, and these values can correspond to the ratio based on pressure or based on temperature.Other values are also can
Can.For example, in a particular embodiment, described two turbines 26,27 can have same or similar reaction ratio；Another
In embodiment, the first stage turbine 26 has the reaction than the second stage turbine 27 than higher reaction ratio.Turbine 26,27 2
Person is configurable to tangential turbine, or is configurable to pressure type turbine both turbine 26,27.
Again referring to Fig. 1, in the illustrated embodiment,（It is multiple）Compressor drum 14',（It is multiple）First order turbine rotor
26' and（It is multiple）Second level turbine rotor 27' is connected to（For example, be rigidly connected to, be integral formed, being attached to or
Allow rotor with axle with the connection of any other type rotated together with identical speed）Turbine shaft 19, the turbine shaft 19
Gearbox module 20 is extended through, it is parallel with engine shaft 16 and radially offset from radial direction（That is, it is not coaxial therewith）.
As that can see in figures 1 and 4,（It is multiple）Compressor drum 14' and（It is multiple）Turbine rotor 26', 27' are outstanding
Arm, i.e., turbine shaft 19 is rotationally supported within（It is multiple）On compressor drum 14' only one sidepiece, and it is supported on
On turbine rotor 26', 27' only one sidepiece.Turbine shaft 19 is by being entirely located in（It is multiple）Compressor drum 14' phase
With on sidepiece, be entirely located in（It is multiple）On first order turbine rotor 26' identical sidepiece and it is entirely located in（It is multiple）The second level
Multiple bearings 64 on turbine rotor 27' identical sidepiece（Such as such as oil lubrication roller bearing and oil lubrication ball bearing
Rolling element bearing, the bearing of journals）To pivotally support.In the illustrated embodiment, bearing 64 is located at（It is multiple）Compressor turns
Between sub- 14' and turbine rotor 26', 27', and it is comprised in gearbox module shell 21, without being arranged on gear-box
The extra bearing in the outside of module 20.Compressor module 24 and the rotary components of turbine module 28 by dynamics be designed as with
Cantilevered fashion rotates, and has the critical deflection mode outside the operating condition of engine.Therefore, compressor module 24 and turbine
Machine module 28 does not include bearing, and is not therefore the part of powered bearing lubricant circulation system 66, the powered bearing lubricant circulation
System 66 is comprised in gearbox module shell 21.This is eliminated on compressor module 24 and turbine module 28 outside setting
Lubricant（Such as oil）Feeding or the needs for cleaning circuit, it can be advantageous to the remainder shifting from hybrid engine component 10
Except compressor module 24 and turbine module 28.
Alternatively, compressor 14 and turbine section 18 can each have itself special axle, such as optimal
Component capabilities.In the case, compressor shaft can also be only by being entirely located in（It is multiple）Compressor drum 14' phase homonymy
In portion（Such as in gearbox module shell 21）Bearing support so that（It is multiple）Compressor drum 14' is propped up in cantilever fashion
Support.（It is multiple）Compressor drum 14' engages with turbine shaft 19 and/or the driving of engine shaft 16, such as passes through gearbox module
20 pinion unit mechanically links compressor shaft and turbine shaft 19 and/or engine shaft 16.
Again referring to Fig. 1, gearbox module 20 is built-up gear tank module 20, including compound gear power train 68 and is included in
One or more annex pinion units in gearbox module shell 21（accessory gear train）Both 70.Turbine
Arbor 19 is mechanically linked to engine shaft 16 by compound gear power train 68 and engaged with the driving of engine shaft 16 so that logical
The mechanical energy for crossing the mechanical energy and engine shaft 16 of the recovery of turbine section 18 is compound.In a particular embodiment, composite teeth
Taking turns power train 68 includes skew gear.In a particular embodiment, the element configuration of compound gear power train 68 allows often to limit
The speed reducing ratio that individual module operates under its Optimum Operation speed.Speed reducing ratio can be accordingly depending upon size of engine and/or other
Factor.In a particular embodiment, speed reducing ratio is approximation 5：1；Other values are also possible.
In a particular embodiment, there is compressor and turbine rotor 14', 26', 27' to allow again on identical axle 19
It is lighter to close pinion unit 68, because compound gear power train is sized to only transmit turbine output in driving compressor
Remaining part after 14.
It should be understood that other types of pinion unit is also possible, particularly, although not being proprietary, for module
Between relative position other configurations.For example, in alternative embodiments, turbine section 18 and/or compressor section 14 can
To be positioned such that its rotary part and the coaxial rotating of engine shaft 16, and planetary gear system can provide engine shaft
Mechanical linkages and driving between 16 and the axle of turbine section 18 and/or compressor section 14 engage.Other configurations are also can
Can.
（It is multiple）Annex pinion unit 70 is by one or more annexes 72 and engine shaft 16 and/or turbine shaft 19
Connection（Mechanically link）.Annex 72 be installed in gearbox module shell 21 with the identical face of compressor module 24, and
It may include but be not limited to one in lower component or any combination：Starter, petrolift, oil pump, cooling medium pump, aircraft liquid
Press pump, aircraft compressor of air conditioner, generator, alternating current generator, permanent magnet alternating current generator.In a particular embodiment, annex
Pinion unit 70 includes skew gear train.Other configurations are also possible, are included but is not limited to, skew and epicyclic train
The combination of system.
With reference to figure 3-4, turbine module 28 is close to nucleus module 22, and positioned at hot side portion（Turbine module 28 and core
Core module 22）With cold side portion（Compressor module 24 and annex 72）Between gearbox module 20 make it possible to define it is relatively small
Area of fire protection（fire zone）, it simplifies the design of aircraft nacelle and fire extinguishing system in a particular embodiment, improves pair
In the fire savety of the remainder of hybrid engine component.In the illustrated embodiment, hybrid engine component 10 includes
The circumferential fire wall 63 for circumferentially extending and extending radially outwardly around gearbox module shell 21.Fire wall 63 is oriented to
So that thermal region or area of fire protection are located on one sidepiece, and annex 72 and compressor module 24 are located at its another sidepiece
On.In a particular embodiment, thermal region includes turbine module 28, blast pipe 30 and nucleus module 22 and the phase of blast pipe 30
Adjacent surface, for example, it is fan-shaped along 75 degree comprising blast pipe 30（quadrant）.In the illustrated embodiment, turbine module
28 and nucleus module 22 be located on a sidepiece of fire wall 63, and annex 72 and compressor module 24 are located at another sidepiece
It is upper --- i.e. the axial location of fire wall 63 is in the axial location of turbine module 28 and nucleus module 22 and annex 72 and compression
Between the axial location of machine module 24.
Additional fire wall is arranged to fuel system 13 and thermal region（Arranged including heat turbine module 28 and turbine
Tracheae 30）Isolation.In the fig. 3 embodiment, two axial fire walls 65,67 extend from circumferential fire wall 63；Axial fire wall
65th, 67 is axially extending along nucleus module 22, and extends radially outwardly.This two axial fire walls 65,67 are circumferentially from one another
It is spaced apart so that fuel system 13 is therebetween；One in fire wall 65 can be positioned so that be located at or around rotation send out
Motivation 12' top dead-centre position.In the illustrated embodiment, axial fire wall 65,67 is oriented to be located at or around 12 respectively
O'clock position（Dieback point）With 4 o'clock position.Additional circumferential fire wall 69 and the first circumferential fire wall 63 are axially spaced, and
And between axial fire wall 65,67, circumferentially about nucleus module 22 part and from nucleus module 22 radially outwardly
Extension.Therefore fuel system 13 is encapsulated in the border limited by fire wall 63,65,67,69, it is by fuel system 13 and whirlpool
Turbine module 28, annex 72 and compressor module 24 separate.
In a particular embodiment, fire wall 63,65,67,69 extends radially outwardly to the position of cabin profile so that cabin
With the boundary fitting limited by fire wall 63,65,67,69, to divide with annex 72, compressor module 24 and turbine module 28
Liftoff encapsulated fuel system 13, and coordinate with the first circumferential fire wall 63, to be separated with annex 72 and compressor module 24
Ground encapsulates turbine module 28 and nucleus module 22.In another embodiment, it is positioned at the additional of the inner radial of cabin profile
Fire wall can be configured to coordinate with fire wall 63,65,67,69, to form the packaging part for including fuel system 13, and
The packaging part includes turbine module 28 and nucleus module 22 independently of cabin, so as to the envelope that provides Yu can be limited by cabin
Smaller packaging part is compared in piece installing.
In a particular embodiment, electric device or annex are not included in turbine module 28, it is reduced or eliminated lets out in fuel
The risk caught fire in the case of leakage in turbine module 28.In addition to the sensor and electric device related to nucleus module 22
Sensor and electric device are entirely located in the cold side portion of gearbox module 20（Temperature is not high enough to catch fire to put in this place）, and
And separated by fire wall 63 with thermal region；Fuel system 13 by fire wall 65,67,69 further with including turbine module
28 and blast pipe 30 thermal region remainder separation, to be further minimized the risk caught fire.
It should be understood that in figure 3, for purposes of clarity, fire wall 63,65,67,69 has been schematically illustrated to be transparent
, to avoid hindering the view of other parts of engine 10, but such illustration does not imply that fire wall 63,65,67,69 needs
It is made of clear material.Fire wall 63,65,67,69 by resisting any material of high temperature according to current certification demand enough
It is made.In a particular embodiment, fire wall 63,65,67,69 is by that can resist 2000 °F of the constant temperature material system of 5 minutes
Into.The example of suitable material is steel, but can use other suitable materials.
With reference to figure 5, hybrid engine component 10 is reflux assembly.Hybrid engine component 10 includes entry conductor 74, tool
Have and surrounding air on the outside of component 10 or around component 10（Such as the environment on the outside of the cabin of receiving unit is empty
Gas）The entrance 76 of connection.Entry conductor 74 includes inertia particle separator 78 in end downstream.In inertia particle separator 78
It is adjacent（immediately）Downstream, entry conductor connect with the first pipeline 80 and the second pipeline 82, the first pipeline 80 and compression
Machine 14 connects, and the second pipeline 82 limits leads with the entrance bypass connected in the outside of component 10 or around component 10 surrounding air
Pipe, such as pass through the exhaust manifolds 84 with hybrid engine component 10（See Fig. 6）Connection.First pipeline 80 is relative to entry conductor
74 sharp turns defined below（For example, by approximately perpendicular to its extension）, it is extended with sufficient angle from entry conductor 74 so that
Heavier particle（Such as ice, sand）Proceed into the second pipeline 82 of downward angle, and urgency of the air along the first pipeline 80
It is curved.The section of the restriction inertia particle separator 78 of entry conductor 74 and the first and second pipelines 80,82 are sized to
Appropriate air velocity is realized to ensure the separation of particle.
Referring still to Fig. 5, during power operation, surrounding air passes through entry conductor on an end of component 10
74 entrance 76 penetrates hybrid engine component 10, and cycles through entry conductor in the first direction across the length of component 10
74.Air reaches compression after having already passed through inertia particle separator 78, being transferred in pipeline 80 and cycling through filter 86
Machine 14.The stream that the modulation of inlet guide vane 88 enters in compressor 14.Air increases pressure by compressor 14 and is directed to hair
Motivation core 12；Although being not shown, the air stream between compressor 14 and engine core 12 can partially or wholly be followed
Ring passes through intercooler.12 further compressed air of engine core.Fuel is injected into engine core 12 and burnt,
And work(is extracted during the expansion cycles of engine core 12.Exhaust from engine core 12 is circulated to turbine
Section 18.Work(is further by turbine（Such as tangential turbine, then pressure type turbine）Extract to drive compressor 14,
And remaining useful work is passed to engine shaft 16 via gearbox module 20.From compressor 14 to turbine section 18
Air/gas circulation is completed along with the substantially opposite direction in the direction of the air circulation in entry conductor 74 so that exhaust gas
Body leaves turbine section 18 near the identical end of entrance 76 with entry conductor 74 of component 10.
In the illustrated embodiment, the part of exhaust stream is used for anti-icing/deicing of the entrance 76 of component 10.Whirlpool
Turbine exhaust connects with first exhaust pipeline 90 and second exhaust pipeline 91, and the first exhaust pipeline 90 connects with exhaust manifolds 84
Logical, the second exhaust pipeline 91 connects with one or more pipelines 92 being located in the lip of entrance 76, it is one or
More pipelines 92 are then gone back（Such as directly）By connected with exhaust manifolds 84 or by with the second pipeline（Entrance bypasses
Conduit）82 connections, with the outside of component 10 or connect around the surrounding air of component 10.Valve 94 can be arranged on second row
The entrance of air pipe 91, to adjust（It is multiple）The evacuating air stream that is circulated in lip pipeline 92 and/or when need not remove
The stream is closed during ice.
Additionally or alternatively, it is anti-icing to pass through the heat exchanger from component 10（Cooler）96（See Fig. 6）It is hot cold
But agent is realized, such as is being circulated to associated heat exchanger by the part for the hot coolant stream for making to leave engine core 12
The helix tube 98 being arranged in the lip of entrance 76 is cycled through before 96.
Referring still to Fig. 5, it can be seen that turbine shaft 19 is parallel to engine shaft 16 and radially offsets from radial direction（I.e., not
It is coaxial therewith）, and from the radial deflection of entry conductor 74 both axle 16,19（That is, it is not coaxial therewith）.In the embodiment shown
In, at least part of axle 16,19 from entry conductor 74 or the longitudinal center axis from whole entry conductor 74 100 are radially inclined
Move.Air stream in entry conductor 74 occurs along the direction for corresponding to or being corresponding generally to central axis 100.In it should be understood that
Mandrel line 100 can be straight line（Straight catheter）Or curve（Bending conduit, such as single curve, S-shaped）.In a particular embodiment,
Central axis 100 is parallel to axle 16,19.Other configurations are also possible, are included but is not limited to, central axis 100 is relative to axle
16th, 19 are extended with the angle of non-zero.There is curved shape in entry conductor 74（Such as）Embodiment in, imaginary line can be limited
It is set to the straight line for the flexural center axis for corresponding more closely to entry conductor 74；The imaginary line can parallel to axle 16,19 or
Person is extended relative to it with the angle of non-zero.
Fig. 6 shows turbine shaft 19, module inlet 76 and entry conductor 74, cycles through hybrid engine for cooling down
The lubricant of the oily or other lubricant of component 10（Such as oil）Heat exchanger 102（For example, so as to lubricate the bearing of axle 16,19 and
Engine core 12（It is multiple）Rotor）And for cool down cycle through engine core 12 housing cooling agent it is cold
But agent（Such as water）The example of the relative angular position of heat exchanger 96.In a particular embodiment, the cloth of hybrid engine component 10
Office is suitable for the compact streamline cabin with minimum aircraft resistance.
Turbine shaft 19 and entry conductor 74 allow compressor and turbine mould relative to the radial deflection of engine shaft 16
Block 24,28, entry conductor 74 and heat exchanger 96,102 can be arranged around engine shaft 16 into clock（clockable）, i.e., to enclose
Arranged around the various angle positions of engine shaft 16, to be adapted to specific aircraft nacelle to design.For example, Fig. 6 configuration can be with
By the way that compressor and turbine module 24,28 are placed as changing closer to cabin exhaust portion, such as more towards component
10 bottom, so as to exhaust manifolds 84 are reduced or minimized and/or be connected to exhaust manifolds 84 gas exhaust piping 90,91 length
Degree.Module inlet 76 and entry conductor 74 can also be changed around the angle position of engine shaft 16, to be adapted to specific fly
Row device cabin design.Cooling agent and lubricant heat exchanger 96,102 can be for example on the sidepiece of nucleus module 22, core moulds
At the top of block 22 or after nucleus module 22, such as it is suitable for the given aircraft associated with hybrid engine component 10
And/or to provide increased the accessing property of heat exchanging device 96,102 and other parts, in order to maintenance, repair and/or more
Change.Annex 72 can be entirely located in identical angle position, and as desired relative to reception hybrid engine component 10
Obtainable space and around nucleus module 22 into clock arrange.In a particular embodiment, whole annexes 72 are positioned at identical
Angle position, which allows whole annexes 72 to access panel by single compartment, to be accessed.
Referring again to Fig. 4, in a particular embodiment, hybrid engine component 10 passes through the installation cage portion including pillar 106
104 are attached to aircraft, and the pillar 106 is connected to two relative sidepiece peaces of the shell 21 for being attached to gearbox module 20
Dress portion 105, and bottom installation portion 105' is also attached to shell 21.In the illustrated embodiment, cage portion 104 is installed by cluster engine
Part 10 is connected to two upper aircraft mount points 108 and two lower aircraft mount point 108'（Such as be arranged on aircraft every
Plate（bulkhead）On）.On each sidepiece, installation cage portion 104 includes be connected to respective side portion installation portion 105 first and the
Two pillars 106, and it is connected to bottom installation portion 105' the 3rd pillar 106''；Pillar 106,106' and 106'' pass through isolation
Portion 103 is connected to installation portion 105,105', and the isolation part 103 may include for example suitable elastomeric material.From identical sidepiece
The first and second pillars 106, the 106' of the extension of installation portion 105 are angled with respect to each other, so as to away from sidepiece installation portion 105
Distance increase and more extend outwardly away from each other.First pillar 106 is configured to be connected to upper aircraft mount point 108 accordingly,
And second and the 3rd pillar 106', 106'' be configured to be connected to lower aircraft mount point 108' accordingly.Arch support portion 107
In the extension of engine 10 times, and be connected to the shell 21 of gearbox module 20, and installation portion 105,105' by with arch branch
The connection of support part 107 is attached to shell 21.Pillar 106,106', 106'' are positioned such that to avoid passing through blast pipe 30.Specific
In embodiment, such configuration avoids making any hot gas leak into turbine module 28 from core-engine blast pipe 30
In, impact in mounting structure（Including isolation part 103, fastener etc.）On, and therefore avoid endangering mounting structure integrality, its
It can be caused by the such leakage impacted on mounting structure.
In the illustrated embodiment, cage portion 104 is installed and installation portion 105,105' is located at area of fire protection（Turbine module
28/ nucleus module 22）Outside.It is located at including pillar 106,106', 106'' installation cage portion 104 and installation portion 105,105'
In " the cold side portion " of gearbox module shell 21, and pass through fire wall 63 and turbine module 28, nucleus module 22 and blast pipe
30 separation.Therefore installation cage portion 104 is completely contained in from the first position at the cold end of component axially extends to gear-box
In the axial space of the second place in module case 21, wherein turbine module 28, nucleus module 22 and blast pipe 30 are located at this
On the outside of axial space.Therefore, pillar 106,106', 106'' be not by the warm of turbine module 28, blast pipe 30 and nucleus module 22
Temperature challenge, this can help improve installation cage portion 104 and its structural intergrity being connected with engine 10.
With reference to figure 7-9 and Figure 10 A-10B, the hybrid engine component 210 according to alternate embodiment is shown, wherein with answering
Close engine pack 10 the same or analogous element of respective element identified by identical reference, and will not herein
In further describe.As shown in figs. 7-8, hybrid engine component 210 is configured to backflow single-shaft engine and including five
（5）Primary module：Nucleus module 22, gearbox module 20, cold section/compressor module 24, hot section/turbine module 28, with
And reduction gearing tank module 220.In hybrid engine component 210, revolved by what the engine shaft 16 of nucleus module 22 drove
It is propeller 208 to turn load.Engine shaft 16 is bonded to propeller 208 by reduction gearing tank module 220.The quilt of nucleus module 12
It is portrayed as including three（3）Rotary engine 12', however, it is understood that any other appropriate number of rotating engine can be used
Machine or other types of explosive motor.
In the illustrated embodiment, reduction gearing tank module 220 includes epicyclic gearbox system；Other configurations are also possible
, include but is not limited to, skew gear-box and double finger offsets pinion units.Although being not shown, additional annex can be by
Mechanically link to and be drivingly engaged to reduction gearing tank module.
With reference to figure 9, in use, surrounding air penetrates hybrid engine component 210 by the entrance 76 of entry conductor 74,
Entry conductor 74 is cycled through, by inertia particle separator 78, changes direction to circulate across filter 86, inlet guide vane
88th, compressor 14, optional intercooler 217（See Figure 10）And engine core 12.Row from engine core 12
Gas is circulated to turbine section 18（It may include foregoing two turbine stages）, wherein work(further extracted with drive
Dynamic compressor.Remaining useful work is passed to engine shaft 16 via gearbox module 20.It can be seen that exhaust stream
Part can be circulated to lip pipeline 92, for entrance 76 lip it is anti-icing, as described above.
Fire wall 63 is from gearbox module shell 21 in area of fire protection（The nucleus module 22 of turbine module 28/）With the He of annex 72
Extend between compressor module 24, as described above.
Hybrid engine component 210 also includes parallel to engine shaft 16 and radially offset from radial direction（That is, it is same not with it
Axle）Turbine shaft 19, two of which axle from along entry conductor 74 all or part of length extension central axis
100 radial deflections（That is, it is not coaxial therewith）.Central axis 100 can parallel to axle 16,19, can be relative to axle 16,19 with
The straight line of the angle extension of non-zero can be bending（For example, single curve, S-shaped）.There is bending in entry conductor 74
In the embodiment of shape, imaginary line can be defined as corresponding more closely to the straight of the flexural center axis of entry conductor 74
Line；The imaginary line can extend parallel to axle 16,19 or relative to it with the angle of non-zero.Turbine shaft 19 and entry conductor
74 allow compressor and turbine module 24,28, entry conductor 74 and heat exchanger relative to the radial deflection of engine shaft 16
96th, 102 can be arranged around engine shaft 16 into clock, i.e., to be arranged around the various angle positions of engine shaft 16, to be adapted to
Specific aircraft nacelle design, as described above.
With reference to figure 10A, hybrid engine component 210 also includes installation cage portion 204, and the installation cage portion 204 includes angulation
Pillar 206, the 206' of degree, its be connected to by arch support portion 107 be attached to two of shell 21 of gearbox module 20 it is relative
Sidepiece installation portion 105, and angled pillar 206'', it is connected to is also attached to by additional arch support portion 207
The bottom installation portion 105' of engine pack 210, the additional 207 and first arch support portion of arch support portion 107 is axially
It is spaced apart, such as is configured to support reduction gearing tank module 220.In this embodiment, cage portion 204 is installed to wrap on each sidepiece
The first and second pillars 206, the 206' for being connected to respective side portion installation portion 105 are included, and is connected to bottom installation portion 105''s
3rd pillar 106''.The first and second pillars 206 for extending from identical sidepiece installation portion 105,206' angulations relative to each other
Degree, more extended outwardly away from each other so as to increase with the distance away from sidepiece installation portion 105.First pillar 206 is configured to connect
To corresponding upper aircraft mount point 108, and second and the 3rd pillar 206', 206'' be configured to be connected to corresponding lower fly
Device mount point 108'.Connecting rod 209 is arranged on each sidepiece to interconnect described two arch support portions 107,207.
As described above, in a particular embodiment, cage portion 204 is installed and installation portion 105 passes through fire wall 63 and whirlpool
Turbine module 28, nucleus module 22 and blast pipe 30 separate.Further, since in this embodiment, the 3rd pillar 206'' and connecting rod
209 extend to the thermal region of fire wall 63（Such as turbine module 28, blast pipe 30 and nucleus module 22 are adjacent with blast pipe 30
Part）Sidepiece at which is positioned, so the element in installation cage portion 204（Pillar 206'' and connecting rod 209）And blast pipe
Between 30, i.e., between the element and thermal region in installation cage portion 204, a fire wall extends transverse to the fire wall 63.Depend on
In its position, this fire wall can be one in foregoing flameproof wall 65,67, or additional fire wall 165（As shown）.
In the illustrated embodiment, cage portion 204 is installed and installation portion 105,105' is located at outside thermal region.Including pillar
206th, 206', 206'' installation cage portion 204 and installation portion 105,105' by fire wall 63 and 165 and turbine module 28,
Blast pipe 30 and nucleus module 22 is adjacent with blast pipe 30 is partially separated.Therefore, pillar 206,206', 206'' be not by turbine
The hot temperature challenge of machine module 28, blast pipe 30 and nucleus module 22, this can help improve installation cage portion 204 and its and engine
The structural intergrity of the connection of component 210.
It should be understood that the configuration in the installation cage portion of engine pack 201 can be with the difference that shows；For example, Fig. 4 installation cage
Portion 104 can be used together with engine pack 210.Similarly, Figure 10 A installation cage portion 204 can be with engine pack 10
It is used together.Other configurations are also possible.For example, installation cage portion 104,204 can include additional pillar.
With reference to figure 11-12, show the hybrid engine component 310 according to alternate embodiment, wherein with hybrid engine
The same or analogous element of respective element of component 10,210 is identified by identical reference, and will not be entered herein
One step describes.Hybrid engine component 310 is configured to backflow single-shaft engine and including four（4）Primary module：Nucleus module
22nd, cold section/compressor module 24, hot section/turbine module 28 and including the first and second submodules or part 320,
320' gearbox module, first and second submodule or part 320,320' coordinate to limit together similar to foregoing
The module of gearbox module 20.Although being not shown, hybrid engine component 310 is configurable to have reduction gearing tank module
Turboprop（turboprop engine）.
In a particular embodiment, hybrid engine component 310 is configured as in addition to its gearbox module 320,320'
It is similar or identical with foregoing hybrid engine component 10 or hybrid engine component 210；It is therefore to be understood that foregoing group
Any element of part 10,210 or the combination of element can be used in component 310.
The Part I 320 of gearbox module include comprising（Such as encapsulate）Compound gear power train（Shown herein as small
Gear（pinion gear））Part I 368 shell 321, and the Part II 320' of gearbox module is included comprising multiple
Close the complementary portion 368' of pinion unit shell 321'.Described two gearbox module shells 321,321' are removably interconnected；
In the illustrated embodiment, shell 321,321' include complementary flange 323,323', and it, which is bolted, links together, and has therebetween
Have and spacer portion 331 is set.However, it is possible to use the connection of any other suitable type, including but not limited to those described above connect.
Turbine shaft 19 extends through the Part II 320' of gearbox module, compressor module 24 and turbine module 28
Rotor be connected to（Such as it is rigidly connected to, is integral formed, being attached to or allowing rotor and axle is with same speed
The connection of any other type rotated together）The turbine shaft 19.The part 368 of compound gear power train, 368' coordinate
Mechanically to link or be drivingly engaged turbine shaft 19 to engine shaft 16.Turbine module 28 and compressor module 24
Rotor be cantilever, and support turbine shaft 19 bearing 64 be comprised in gearbox module Part II 320' shell
In 321', without the additional bearing being arranged on the outside of gearbox module.Alternatively, turbine module 28 and compressor mould
Block 24 can each have the special axle of their own.Compressor module 24 and turbine module 28 do not include bearing, and therefore
It is not the part of powered bearing lubricant circulation system, the powered bearing lubricant circulation system is comprised in second gear tank module shell
In 321'.
Compressor module shell 25 is located at gearbox module shell 321,321' outsides, and is installed in second gear tank module
On shell 321' face（For example, the connection by any suitable type（Including but not limited to those described above connects）It is removably mutual
Even）.Turbine module shell 29 also is located at gearbox module shell 321,321' outsides, and is installed in second gear tank module shell
On the 321' face relative with the face of reception compressor module shell 25（For example, the connection by any suitable type（Including but not
It is limited to those described above connection）It is removably interconnected）.
The Part I 320 of gearbox module includes one or more included in first gear tank module shell 321
Annex pinion unit（It is not shown）.Annex（It is not shown）Gearbox module 320,320' with the identical of compressor module 25
It is mounted in engagement with sidepiece on the face of first gear tank module shell 321.
Gearbox module shell 321, the 321' of separation can allow turbine module 28, compressor module 24 and gearbox module
Part II 320' separated with the remainder of engine 310, while keep interconnected amongst one another, to limit " turbomachinery mould
Block ", it independently can be replaced or maintain with the remainder of engine 310.
In a particular embodiment, the gearbox module shell 321 of separation, 321' allow second adjacent with turbine module 28
Shell 321' is made up of material of the material than first shell 321 more resistant to heat, with the single tooth completely made of the material of first shell 321
Roller box module case is on the contrary, it can help to minimize cooling requirement and/or Thermal protection demand.In a particular embodiment, first shell
321 is made of aluminum, and second shell 321' is formed from steel.
Although being not shown, engine 310 includes installation portion, for being engaged with mounting structure, such as foregoing installation
Cage portion 104,204.In a particular embodiment, installation portion is connected to first gear tank module shell 321.
Although the example of hybrid engine component 10,210,310 has been shown as turbine wheel shaft and turboprop
Component, it should be appreciated that hybrid engine component can be designed to other purposes, including but not limited to as auxiliary power list
Member.
Foregoing description be intended only it is exemplary, and it would be recognized by those skilled in the art that without departing from disclosed hair
In the case of bright scope, the embodiment of description can be made a change.By looking back the disclosure, fall within the scope of the present invention
Modification will be apparent to those skilled in the art, and it is such modification be intended to fall under appended claims
It is interior.
Claims (20)
1. a kind of hybrid engine component, including：
Engine core, including at least one explosive motor engaged with engine shaft driving；
Turbine section, there is the entrance by least one blast pipe and the communication of the engine core, institute
Stating turbine section includes being connected at least one turbine rotor of turbine shaft, the turbine shaft and the engine shaft
Driving engagement；
Compressor, has the outlet being in fluid communication with the entrance of the engine core, and the compressor includes and the turbine
At least one compressor drum of at least one driving engagement in arbor and the engine shaft；
It is connected to the shell of the turbine section, the compressor and the engine core；And
Cage portion is installed, for the hybrid engine component to be attached into aircraft, the installation cage portion includes being connected to installation
Multiple pillars in portion, the installation portion are attached to the shell on the outside of thermal region, and the thermal region includes the turbine portion
The part adjacent with least one blast pipe of section, at least one blast pipe and the engine core, it is described
Pillar is separated by least one fire wall with the thermal region.
2. hybrid engine component as claimed in claim 1, wherein, all the pillar extends outwardly away from institute from the installation portion
State turbine section and engine core.
3. hybrid engine component as claimed in claim 1, wherein, the shell is the shell of gearbox module, the compressor
On a sidepiece of the shell, the turbine section and engine core are located on the relative sidepiece of the shell, institute
Turbine shaft is stated to extend through the gearbox module and by the pinion unit of the gearbox module with described start
Arbor driving engagement, at least one compressor drum are connected to the turbine shaft.
4. hybrid engine component as claimed in claim 1, wherein, at least one fire wall is included from the shell radially
The fire wall to stretch out, the fire wall is between the installation portion and the thermal region.
5. hybrid engine component as claimed in claim 1, wherein, the compressor is located on a sidepiece of the shell,
The thermal region is located on the relative sidepiece of the shell, and at least one fire wall is included in the installation portion and the heat
The first fire wall to be extended radially outwardly between region from the shell, the multiple pillar be included in first fire wall with
At least one pillar extended on the thermal region identical sidepiece, at least one fire wall are included in described at least one
The second fire wall extended laterally between pillar and the thermal region from first fire wall.
6. hybrid engine component as claimed in claim 1, wherein, the installation portion includes being connected to the relative of the shell
Two sidepiece installation portions of sidepiece, and bottom installation portion, and the multiple pillar is in each sidepiece in the installation cage portion
On include：
From two pillars of a corresponding extension for described two sidepiece installation portions, for being connected to two different aircraft
Mount point, described two pillars are angled with respect to each other, so as to away from the sidepiece installation portion distance increase and more
Extend outwardly away from each other, and
From the additional pillar of bottom installation portion extension, for being connected in described two different aircraft mount points
One.
7. hybrid engine component as claimed in claim 1, wherein, each bag at least one explosive motor
Include and seal and rotatably receive the rotor in respective inner intracavitary, to provide the variable-volume in the respective inner chamber
Rotation chamber, the rotor have separate it is described rotation chamber and be mounted for the respective inner intracavitary bias return
Three top sections turned, the respective inner chamber have the epitrochoid shape with two lug bosses.
8. hybrid engine component as claimed in claim 1, wherein, the turbine section includes having and the engine
First stage turbine of the entrance of the communication of core, and with the outlet fluid with first stage turbine
Second stage turbine of the entrance of connection.
9. hybrid engine component as claimed in claim 7, wherein, the first order turbine configurations are to be up to value
0.2 based on the reaction of pressure than tangential turbine, second stage turbine have than first stage turbine
Reaction is than higher reaction ratio.
10. a kind of hybrid engine component, including：
Engine core, including at least one explosive motor engaged with engine shaft driving；
Gearbox module, including the gearbox module shell comprising at least one pinion unit；
Turbine section, in the outside of the gearbox module shell, the turbine section, which has, passes through at least one blast pipe
With the entrance of the communication of the engine core, the turbine section includes being connected at least the one of turbine shaft
Individual turbine rotor, one at least one pinion unit that the turbine shaft passes through the gearbox module with
The engine shaft driving engagement；
Compressor, in the outside of the gearbox module shell, the compressor has the entrance fluid with the engine core
The outlet of connection, the compressor include engaged with least one driving in the turbine shaft and the engine shaft to
A few compressor drum；
Wherein, the turbine section and the engine core are located on the identical sidepiece of the gearbox module shell, and
And the compressor is located on the relative sidepiece of the gearbox module shell；And
Cage portion, for the hybrid engine component to be attached into aircraft and is connected to the shell, the installation cage are installed
Portion is kept completely separate by least one fire wall and the turbine section and at least one blast pipe.
11. hybrid engine component as claimed in claim 10, wherein, the installation cage portion is completely contained in following axially empty
In, the axial space is axially extended in the gearbox module shell from the first position at the cold end in the component
On the second place, the turbine section and engine core are located at the outside of the axial space.
12. hybrid engine component as claimed in claim 10, wherein, at least one fire wall is included from the gear
The first fire wall that tank module shell extends radially outwardly, the turbine section and engine core are located at first fire wall
Identical sidepiece on, and the compressor and the installation cage portion are located on the relative sidepiece of first fire wall,
The installation cage portion be included in first fire wall with prolonging in the turbine section and engine core identical sidepiece
At least one pillar stretched, at least one fire wall are included between at least one pillar and the turbine section
And second extended laterally between at least one pillar and at least one blast pipe from first fire wall
Fire wall.
13. hybrid engine component as claimed in claim 11, wherein, the installation cage portion is included all from the gear-box
Multiple pillars that module case extends towards the cold end of the component.
14. hybrid engine component as claimed in claim 10, wherein, the installation cage portion includes on each of which sidepiece：
Two pillars extended from the identical sidepiece installation portion for being connected to the gearhousing, it is different for being connected to two
Aircraft mount point, described two pillars are angled with respect to each other, so as to the distance increase away from the sidepiece installation portion
And more extend outwardly away from each other, and
From bottom installation portion extension additional pillar, for be connected in described two different aircraft mount points one
It is individual.
15. hybrid engine component as claimed in claim 14, wherein, each sidepiece installation portion passes through in the gear-box mould
The additional supporting part extended under block is connected to the gearhousing.
16. hybrid engine component as claimed in claim 10, wherein, each at least one explosive motor
Including sealing and rotatably receiving the rotor in respective inner intracavitary, so as to provide in the respective inner chamber can variant
Long-pending rotation chamber, the rotor, which has, to be separated the rotation chamber and is mounted for the bias of the respective inner intracavitary
Three top sections of revolution, the respective inner chamber have the epitrochoid shape with two lug bosses.
17. hybrid engine component as claimed in claim 10, wherein, at least one compressor drum is connected to described
Turbine shaft, the turbine shaft extend through the gearbox module.
18. hybrid engine component as claimed in claim 10, wherein, the turbine section includes having starts with described
First stage turbine of the entrance of the communication of machine core, and with the outlet stream with first stage turbine
Second stage turbine of the entrance of body connection.
19. hybrid engine component as claimed in claim 18, wherein, the first order turbine configurations are most with value
For 0.2 based on the reaction of pressure than tangential turbine, second stage turbine has than first stage turbine
Reaction than higher reaction ratio.
20. hybrid engine component as claimed in claim 10, wherein, the turbine shaft is parallel with the engine shaft
And radial deflection each other.
Applications Claiming Priority (7)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US201562118914P | 2015-02-20 | 2015-02-20 | |
| US62/118914 | 2015-02-20 | ||
| US14/864,124 US10533500B2 (en) | 2015-02-20 | 2015-09-24 | Compound engine assembly with mount cage |
| US14/864124 | 2015-09-24 | ||
| US15/047,362 US10533492B2 (en) | 2015-02-20 | 2016-02-18 | Compound engine assembly with mount cage |
| US15/047362 | 2016-02-18 | ||
| PCT/CA2016/050160 WO2016131148A1 (en) | 2015-02-20 | 2016-02-19 | Compound engine assembly with mount cage |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN107429614A true CN107429614A (en) | 2017-12-01 |
| CN107429614B CN107429614B (en) | 2020-03-10 |
Family
ID=60423850
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201680023261.2A Active CN107429614B (en) | 2015-02-20 | 2016-02-19 | Composite engine assembly with mounting cage |
Country Status (2)
| Country | Link |
|---|---|
| CN (1) | CN107429614B (en) |
| ES (1) | ES2829959T3 (en) |
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| CN108121252A (en) * | 2017-12-15 | 2018-06-05 | 石家庄飞机工业有限责任公司 | A kind of aircraft electrical control system based on full authority digital electronic control system |
| CN108298095A (en) * | 2017-12-15 | 2018-07-20 | 石家庄飞机工业有限责任公司 | A kind of Small General Aircraft boat coal piston engine power plants |
| CN111942599A (en) * | 2019-05-17 | 2020-11-17 | 普拉特 - 惠特尼加拿大公司 | Mounting arrangement for torque rolling vibration isolation |
| CN112814787A (en) * | 2019-11-18 | 2021-05-18 | 上海尚实能源科技有限公司 | Gas turbine engine accessory system integrated unit |
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| US2249194A (en) * | 1938-06-10 | 1941-07-15 | Rolland S Trott | Mounting for airplane engines |
| DE742824C (en) * | 1940-12-28 | 1943-12-11 | Auto Union A G | Multi-cylinder internal combustion engine, especially for aircraft |
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| CN108121252A (en) * | 2017-12-15 | 2018-06-05 | 石家庄飞机工业有限责任公司 | A kind of aircraft electrical control system based on full authority digital electronic control system |
| CN108298095A (en) * | 2017-12-15 | 2018-07-20 | 石家庄飞机工业有限责任公司 | A kind of Small General Aircraft boat coal piston engine power plants |
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| CN112814787B (en) * | 2019-11-18 | 2023-11-24 | 上海尚实航空发动机股份有限公司 | Gas turbine engine accessory system integrated unit |
Also Published As
| Publication number | Publication date |
|---|---|
| ES2829959T3 (en) | 2021-06-02 |
| CN107429614B (en) | 2020-03-10 |
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