CN110901944B - Helicopter engine installation design method - Google Patents
Helicopter engine installation design method Download PDFInfo
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- CN110901944B CN110901944B CN201911227775.1A CN201911227775A CN110901944B CN 110901944 B CN110901944 B CN 110901944B CN 201911227775 A CN201911227775 A CN 201911227775A CN 110901944 B CN110901944 B CN 110901944B
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- 238000009434 installation Methods 0.000 title claims abstract description 32
- 238000000034 method Methods 0.000 title claims abstract description 14
- 238000006243 chemical reaction Methods 0.000 claims description 22
- 239000003638 chemical reducing agent Substances 0.000 claims description 12
- 230000005540 biological transmission Effects 0.000 claims description 3
- 230000003068 static effect Effects 0.000 claims description 3
- RZVHIXYEVGDQDX-UHFFFAOYSA-N 9,10-anthraquinone Chemical compound C1=CC=C2C(=O)C3=CC=CC=C3C(=O)C2=C1 RZVHIXYEVGDQDX-UHFFFAOYSA-N 0.000 claims 1
- 238000005457 optimization Methods 0.000 abstract description 4
- 239000013585 weight reducing agent Substances 0.000 abstract description 3
- 238000004364 calculation method Methods 0.000 description 2
- 230000008878 coupling Effects 0.000 description 2
- 238000010168 coupling process Methods 0.000 description 2
- 238000005859 coupling reaction Methods 0.000 description 2
- 238000012938 design process Methods 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 238000013178 mathematical model Methods 0.000 description 1
- 238000004088 simulation Methods 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
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Classifications
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
- B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
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- B64D27/40—
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- 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
- Y02T50/00—Aeronautics or air transport
- Y02T50/40—Weight reduction
-
- 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
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Abstract
The invention belongs to the technical field of helicopter strength design, and discloses a helicopter engine installation design method, which comprises the following steps: s1, determining the installation load of an engine; s2, deploying a balance mode of the installation load of the engine; s3, carrying out load distribution on the installation load of the engine according to the balance mode of the installation load of the engine; s4, designing an in-plane bracket for mounting the engine according to the load distribution; the quick design of the engine mounting structure can be realized, and the weight reduction optimization design of the mounting structure can be realized.
Description
Technical Field
The invention belongs to the technical field of helicopter strength design, and particularly relates to a helicopter engine installation design method.
Background
The engine is used as the only source of power of the helicopter, and the mounting device of the engine needs to have double requirements of high reliability and low weight cost. The traditional design method is usually non-forward development, the design process needs to be continuously iterated by design and calculation, the design workload is large, and the design period is long.
Disclosure of Invention
In view of the above problems in the background art, an object of the present invention is to provide a method for designing a helicopter engine installation, which can achieve a rapid design of an engine installation structure and a weight reduction optimization design of the engine installation structure.
In order to achieve the purpose, the invention adopts the following technical scheme to realize.
A helicopter engine mount design method, said method comprising:
s1, determining the installation load of an engine;
s2, deploying a balance mode of the installation load of the engine;
s3, carrying out load distribution on the installation load of the engine according to the balance mode of the installation load of the engine;
and S4, designing an in-plane bracket for mounting the engine according to the load distribution.
The technical scheme of the invention has the characteristics and further improvements that:
(1) S1 specifically comprises the following steps:
defining an engine mounting coordinate system OXYZ as a Cartesian coordinate system, wherein the X direction is the direction of an engine output shaft, and the Z direction is vertically upward;
thus, the mounting load of the engine is determined as: inertial loads F of engine in X, Y and Z directions generated by airplane maneuvering x 、F y 、F z (ii) a Reaction torque M generated by engine output torque x (ii) a Gyroscopic moments M generated by coupling the pitching and yawing movements of an aircraft with the rapid rotation of the rotor of an engine y 、M z 。
(2) A power output shaft sleeve is arranged between the engine and the main speed reducer, and the power output shaft sleeve and the main speed reducer are provided with a mounting point A;
s2 specifically comprises the following steps:
inertial load F in X direction x The inertia load F in the Y direction is balanced by the mounting point A of the power output shaft sleeve and the main speed reducer y By Y-reaction F of the support B in the YOZ plane perpendicular to the power take-off shaft By Balanced, Z-direction inertial loads F z Z-direction reaction force F of bracket B in YOZ plane perpendicular to power output shaft Bz Balancing;
reaction torque M x Load balancing through two mounting points of the bracket B in the YOZ plane; moment M of gyro y Z-direction reaction force F through mounting point A Az And Z-direction reaction force F of bracket B Bz Resulting in a moment balance, gyro moment M z Y-direction reaction force F through mounting point a Ay Equilibrium and Y-reaction force F of support B By The resulting moment is balanced.
(3) S2 further comprises: the following requirements are satisfied when the balance mode of the installation load of the engine is deployed:
the power output shaft of the engine and the input end of the main reducer are concentric shafts;
an axial clearance is provided between the in-plane bracket B and the engine or the in-plane bracket B can move axially to compensate the thermal expansion of the engine.
(4) S3 specifically comprises the following steps:
establishing a load balance equation and calculating the safetyLoading F of loading point A in X, Y and Z directions Ax 、F Ay 、F Az Load of bracket B in dough
Load of the in-plane support B->
Component F comprising two mounting points B1 、F B2 ；
Establishing a load balance equation and a geometric equation, and solving F by taking alpha and beta as random variables B1 、F B2 Wherein alpha and beta are respectively F B1 、F B2 Is inclined to the Z direction.
(5) S4 specifically comprises the following steps:
the mounting point A is a spherical hinge which only transmits load;
two main force transmission path directions and component forces F of two mounting points of in-plane bracket B B1 、F B2 The load directions of the two are consistent;
taking the stress sigma of the in-plane bracket B as a variable according to the component force F of the two mounting points B1 、F B2 And the requirements of static strength, stability and fatigue life of the in-plane bracket B, and the section size of the in-plane bracket B is defined.
The design method for mounting the engine can shorten the design period, reduce the design cost and improve the design quality.
Drawings
FIG. 1 is a schematic view of a macro-stress analysis of an engine according to an embodiment of the present invention;
fig. 2 is a schematic view of a force analysis of an in-plane stent according to an embodiment of the present invention.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be obtained by a person skilled in the art without making any creative effort based on the embodiments in the present invention, belong to the protection scope of the present invention.
The embodiment of the invention provides a helicopter engine installation design method, which comprises the following steps:
s1, determining the installation load of an engine;
s2, deploying a balance mode of the installation load of the engine;
s3, carrying out load distribution on the installation load of the engine according to the balance mode of the installation load of the engine;
and S4, designing an in-plane bracket for mounting the engine according to the load distribution.
The technical scheme of the invention has the characteristics and further improvements that:
further, S1 specifically includes:
defining an engine mounting coordinate system OXYZ as a Cartesian coordinate system, wherein the X direction is the direction of an engine output shaft, and the Z direction is vertically upward;
thus, the mounting load of the engine is determined as: inertial loads F of engine in X, Y and Z directions generated by airplane maneuvering x 、F y 、F z (ii) a Reaction torque M generated by engine output torque x (ii) a Gyroscopic moments M generated by coupling the pitching and yawing movements of an aircraft with the rapid rotation of the rotor of an engine y 、M z 。
A power output shaft sleeve is arranged between the engine and the main speed reducer, and the power output shaft sleeve and the main speed reducer are provided with a mounting point A;
s2 specifically comprises the following steps:
as shown in fig. 1, the inertial load F in the X direction x The inertia load F in the Y direction is balanced by the mounting point A of the power output shaft sleeve and the main speed reducer y By Y-reaction F of the support B in the YOZ plane perpendicular to the power take-off shaft By Balanced, Z-direction inertial loads F z Z-direction reaction force F of bracket B in YOZ plane perpendicular to power output shaft Bz Balancing;
reaction torque M x Through two mounting points of the support B in the YOZ planeLoad balancing; moment M of gyro y Z-direction reaction force F through mounting point A Az And Z-direction reaction force F of bracket B Bz Resulting in a moment balance, gyro moment M z Y-direction reaction force F through mounting point a Ay Equilibrium and Y-reaction force F of support B By The resulting moment is balanced.
S2 further comprises: the following requirements are satisfied when the balance mode of the installation load of the engine is deployed:
the power output shaft of the engine and the input end of the main speed reducer are concentric shafts;
an axial clearance is provided between the in-plane bracket B and the engine or the in-plane bracket B can move axially to compensate the thermal expansion of the engine.
S3 specifically comprises the following steps:
establishing a load balance equation, and calculating the load F of the mounting point A in the X direction, the Y direction and the Z direction Ax 、F Ay 、F Az Load of bracket B in dough
Load of the in-plane support B->
Component F comprising two mounting points B1 、F B2 ；
Establishing a load balance equation and a geometric equation, and solving F by taking alpha and beta as random variables B1 、F B2 Wherein alpha and beta are respectively F B1 、F B2 Is inclined to the Z direction.
S4 specifically comprises the following steps:
the mounting point A is a spherical hinge which only transmits load;
as shown in FIG. 2, the two main force transmission path directions and the force components F of the two mounting points of the in-plane bracket B B1 、F B2 The load directions of the two are consistent;
taking the stress sigma of the in-plane bracket B as a variable according to the component force F of the two mounting points B1 、F B2 And the static strength, stability, fatigue life requirements of the in-plane stent B, defining a planeThe cross-sectional dimension of the inner support B.
And finally, establishing a finite element model for engine installation through finite element simulation calculation, and checking whether the strength and the like of the installation structure meet the installation requirements.
The design method of the engine support is based on the optimization solution of the mathematical model, can realize the rapid design of the engine mounting structure, and can realize the weight reduction optimization design of the mounting structure.
The foregoing is illustrative of the present invention and is not to be construed as limiting thereof. The scope of the present invention is not limited thereto, and any changes or substitutions that can be easily made by those skilled in the art within the technical scope of the present invention will be covered by the scope of the present invention. The protection scope of the present invention shall be subject to the protection scope of the claims.
Claims (2)
1. A helicopter engine mount design method, said method comprising:
s1, determining the installation load of an engine; s1 specifically comprises the following steps:
defining an engine mounting coordinate system OXYZ as a Cartesian coordinate system, wherein the X direction is the direction of an engine output shaft, and the Z direction is vertically upward;
thus, the mounting load of the engine is determined as: inertial loads of engine in X, Y and Z directions generated by airplane maneuvering
、/>
、/>
(ii) a Counter-torque generated by the output torque of the engine>
(ii) a Aircraft with a flight control deviceCoupled with the rapid rotation of the motor rotor, generates a gyroscopic moment pick>
、/>
；
S2, deploying a balance mode of the installation load of the engine; a power output shaft sleeve is arranged between the engine and the main speed reducer, and the power output shaft sleeve and the main speed reducer are provided with mounting points
；
S2 specifically comprises the following steps:
inertial load in X direction
Through the counterforce of a mounting point A of the power output shaft sleeve and the main speed reducer>
Counterbalanced, inertial load in the Y-direction->
By means of a holder in the YOZ plane perpendicular to the power take-off shaft>
Is greater than or equal to the Y-direction reaction force>
Balanced, Z-direction inertial load->
Through a bracket in the YOZ plane perpendicular to the power take-off shaft>
Is greater than or equal to the Z-direction reaction force>
Balancing;
reaction torque
Load balancing through two mounting points of the bracket B in the YOZ plane; peg-top moment->
By means of a mounting point->
Is greater than or equal to the Z-direction reaction force>
And the Z-direction reaction force of the support B->
Resulting moment balance, gyro moment>
By means of mounting points>
In the Y direction of (B) < CHEM >>
And the Y-direction reaction force of the support B->
The formed moment balance;
the following requirements are satisfied when the balance mode of the installation load of the engine is deployed:
the power output shaft of the engine and the input end of the main speed reducer are concentric shafts;
an axial clearance is formed between the inner support B and the engine or the inner support B can axially move so as to compensate the thermal expansion of the engine;
s3, carrying out load distribution on the installation load of the engine according to the balance mode of the installation load of the engine; s3 specifically comprises the following steps:
establishing a load balance equation, and calculating the loads of the installation point A in the X direction, the Y direction and the Z direction
、/>
、/>
And in-plane support>
Is greater or less than>
(ii) a In-plane support>
Is greater or less than>
Force component comprising two mounting points->
、/>
；
Establishing load balance equations and geometric equations to
、/>
Is a random variable, evaluated>
、/>
In which->
、
Are respectively based on>
、/>
The included angle between the direction of (a) and the Z direction;
and S4, designing an in-plane bracket for mounting the engine according to the load distribution.
2. A helicopter engine installation design method according to claim 1, characterized in that S4 specifically comprises:
the mounting point A is a spherical hinge which only transmits load;
in-plane stent
Is based on the force component of the two main force transmission path directions and the two mounting points->
、/>
The load directions of the two are consistent;
in-plane support
Is greater than or equal to>
Based on the force component of two mounting points>
、/>
And the requirements of static strength, stability and fatigue life of the in-plane bracket B, define the in-plane bracket->
Cross-sectional dimension (d). />
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
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| CN201911227775.1A CN110901944B (en) | 2019-12-04 | 2019-12-04 | Helicopter engine installation design method |
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| Application Number | Priority Date | Filing Date | Title |
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| CN201911227775.1A CN110901944B (en) | 2019-12-04 | 2019-12-04 | Helicopter engine installation design method |
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| CN110901944B true CN110901944B (en) | 2023-03-31 |
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| CN112052531B (en) * | 2020-09-25 | 2022-12-30 | 中国直升机设计研究所 | Helicopter harpoon load calculation method |
| CN112461525B (en) * | 2020-11-20 | 2022-11-04 | 中国直升机设计研究所 | Unmanned helicopter engine mounting bracket test device |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| US10899462B2 (en) * | 2015-09-04 | 2021-01-26 | Lord Corporation | Anti-torque aft-mounting systems, devices, and methods for turboprop/turboshaft engines |
| CN105270638A (en) * | 2015-11-17 | 2016-01-27 | 江西洪都航空工业集团有限责任公司 | Aeroengine mounting device |
| US20170259929A1 (en) * | 2016-03-10 | 2017-09-14 | General Electric Company | Method and system for mounting an aircraft engine |
| CN109562841B (en) * | 2016-08-08 | 2022-03-18 | 洛德公司 | Mounting system, device and method for aircraft engine |
| CN107563039A (en) * | 2017-08-28 | 2018-01-09 | 中国航空工业集团公司沈阳飞机设计研究所 | A kind of intensity layout design method of aircraft engine installation system |
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