CN109268168A - Height pushes away the small-size turbojet engine of ratio - Google Patents
Height pushes away the small-size turbojet engine of ratio Download PDFInfo
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- CN109268168A CN109268168A CN201811421442.8A CN201811421442A CN109268168A CN 109268168 A CN109268168 A CN 109268168A CN 201811421442 A CN201811421442 A CN 201811421442A CN 109268168 A CN109268168 A CN 109268168A
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- 238000002485 combustion reaction Methods 0.000 claims abstract description 50
- 239000000446 fuel Substances 0.000 claims abstract description 16
- 238000003466 welding Methods 0.000 claims description 2
- 239000000295 fuel oil Substances 0.000 abstract description 20
- 230000000149 penetrating effect Effects 0.000 abstract description 3
- 230000006835 compression Effects 0.000 description 11
- 238000007906 compression Methods 0.000 description 11
- 239000007789 gas Substances 0.000 description 9
- 238000010586 diagram Methods 0.000 description 8
- 230000000694 effects Effects 0.000 description 5
- 239000012530 fluid Substances 0.000 description 5
- 238000000034 method Methods 0.000 description 4
- 239000002737 fuel gas Substances 0.000 description 3
- 238000002347 injection Methods 0.000 description 3
- 239000007924 injection Substances 0.000 description 3
- 239000000463 material Substances 0.000 description 3
- 239000000243 solution Substances 0.000 description 3
- 239000007921 spray Substances 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 2
- 230000008859 change Effects 0.000 description 2
- 230000008569 process Effects 0.000 description 2
- 230000001133 acceleration Effects 0.000 description 1
- 230000009471 action Effects 0.000 description 1
- 238000004458 analytical method Methods 0.000 description 1
- 230000001914 calming effect Effects 0.000 description 1
- 239000000567 combustion gas Substances 0.000 description 1
- 230000006837 decompression Effects 0.000 description 1
- 230000001419 dependent effect Effects 0.000 description 1
- 238000004401 flow injection analysis Methods 0.000 description 1
- 230000005484 gravity Effects 0.000 description 1
- WJRBRSLFGCUECM-UHFFFAOYSA-N hydantoin Chemical compound O=C1CNC(=O)N1 WJRBRSLFGCUECM-UHFFFAOYSA-N 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 239000000203 mixture Substances 0.000 description 1
- 238000003032 molecular docking Methods 0.000 description 1
- 230000008450 motivation Effects 0.000 description 1
- 239000003921 oil Substances 0.000 description 1
- 238000005457 optimization Methods 0.000 description 1
- 238000005476 soldering Methods 0.000 description 1
- 239000007787 solid Substances 0.000 description 1
- 230000007704 transition Effects 0.000 description 1
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/28—Rotors specially for elastic fluids for centrifugal or helico-centrifugal pumps for radial-flow or helico-centrifugal pumps
- F04D29/284—Rotors specially for elastic fluids for centrifugal or helico-centrifugal pumps for radial-flow or helico-centrifugal pumps for compressors
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/28—Rotors specially for elastic fluids for centrifugal or helico-centrifugal pumps for radial-flow or helico-centrifugal pumps
- F04D29/30—Vanes
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The present invention provides a kind of high small-size turbojet engines for pushing away ratio, are related to aero-engine field, solve the low technical problem of existing turbojet engine thrust ratio；The device includes air intake duct, compressor, combustion chamber, turbine and jet pipe, includes spaced primary blades and splitterr vanes on the rotor of compressor, primary blades are distorted shape, and the distance between primary blades are gradually increased by central axis to the surrounding of rotor；Compressor rotor is connect by turbine wheel shaft with turbine rotor；Combustion chamber is toroidal combustion chamber comprising the end wall of burner inner liner and fuel nozzle, burner inner liner is opposite with compressor and it is arc, and fuel nozzle is set on end wall, and multiple air inlets are provided on the side wall of burner inner liner；It is expanding channel between the rotor blade of compressor of the present invention, fuel oil enters from the penetrating of burner inner liner rear, air from burner inner liner side wall, and burning sufficiently, meets target drone and carries out the requirement that high maneuver uses, and has the advantages that light-weight, starting is simple, thrust ratio is big.
Description
Technical field
The present invention relates to aero-engine technology fields, more particularly, to a kind of high small-size turbojet engine for pushing away ratio.
Background technique
Aero-engine is aircraft " heart ", largely determines the development level of aeronautical technology.Make a general survey of boat
The development of empty space technology, each time major progress all be unable to do without the development of aero engine technology.Due to aero-engine
The particularity of application, it is undoubtedly the most advanced level for having embodied a concentrated reflection of modern dynamic power machine, and represents engine of new generation
The developing direction of tool.The main body of current aerospace engine is still gas-turbine unit.Gas-turbine unit is with its power
Greatly, outstanding advantages small in size are widely used in aviation power.
The one-hundred-year history of aero-engine is broadly divided into two periods.Head of first period from 1903 Nian Laite brother
Secondary flight starts until World War II terminates.Within this period, piston-mode motor has been ruled 40 years or so, the
Two periods terminate so far from World War II.Over more than 70 years, aero gas turbine engine is instead of piston engine
Machine has started jet age.Currently, aero gas turbine engine occupies the leading position of aviation power.The Hui Teer of Britain is public
The glycolylurea company difficult to understand of department and Germany has succeeded in developing centrifugal-flow jet engine in July nineteen thirty-seven and nineteen thirty-seven September respectively
WU and HeS3B, HeS3B are the jet planes that first makes a successful trial flight, and have started jet-propulsion new era and aviation industry
New era.The practical turbojet of First is the You Mo -004 of Germany in the world.But it is domestic at present available
The small-size turbojet engine type kind that uses is simultaneously few.
The thrust ratio of engine is the ratio of motor power He weight itself, mitigates the weight of engine, improves it and push away
Power can be improved the thrust ratio of engine.In general, the temperature when air-flow comes out from combustion chamber is higher, the energy of input is got over
Greatly, the thrust of engine is also bigger.But due to the limitation of turbine material etc., can only achieve 1650K or so, in order to increase
The thrust of engine, existing means are to allow not sufficiently combusted combustion along with after-burner sprays into fuel oil after turbine
Gas is mixed with the fuel oil of penetrating to burn again, since component, temperature can make to start up to 2000K without spin in afterbunring room
The thrust of machine increases.But its disadvantage is exactly that oil consumption sharply increases, while excessively high temperature also influences the service life of engine.
The compression ratio of engine is the ratio of the pressure before the pressure and compression after air is compressed by the compressor.Combustion chamber
It when the compressor operating of front end, needs to be pressurized the air of sucking repeatedly, to mix with fuel oil later and generate height
High-speed flow injection is pressed, thrust is generated.Compression ratio is higher, more advantageous for the generation of thrust, for reducing fuel consumption
It is advantageous.
Applicants have discovered that at least there is following technical problem in the prior art: existing small-size turbojet engine thrust ratio phase
To lower, the compressor part compression in engine is relatively low, and fuel burning is insufficient in combustion chamber, so that turbojet engine work
Make inefficient, performance is poor.
Summary of the invention
It is existing in the prior art to solve the purpose of the present invention is to provide a kind of high small-size turbojet engine for pushing away ratio
Turbojet engine thrust ratio is low, the undesirable technical problem of performance；Optimization technique in many technical solutions provided by the invention
Many technical effects elaboration as detailed below that scheme can be generated.
To achieve the above object, the present invention provides following technical schemes:
Height provided by the invention pushes away the small-size turbojet engine of ratio, including air intake duct, compressor, combustion chamber, turbine and tail
Jet pipe, includes spaced primary blades and splitterr vanes on the rotor of the compressor, and the primary blades are distorted shape, and institute
The distance between primary blades are stated to be gradually increased by central axis to the surrounding of rotor；Compressor rotor is turned by turbine wheel shaft and turbine
Son connection；
The combustion chamber be toroidal combustion chamber comprising burner inner liner and fuel nozzle, the end wall of the burner inner liner with it is described
Compressor is opposite and it is arc, and the fuel nozzle is set on the end wall, is provided on the side wall of the burner inner liner
Multiple air inlets.
Preferably, the air intake duct is Subsonic inlet, and the inside of the air intake duct is by entrance to the compressor
Place gradually tapers up；And the air intake duct is internally provided with rectification supporting plate and fiaring cone.
Preferably, the compressor further includes radial rectifier and axial rectifier, the radial direction rectifier and the axis
It is set between the compressor rotor and the combustion chamber to rectifier, is successively passed through by the air-flow of the compressor rotor
The radial direction rectifier and the axial rectifier, directional velocity change into axial direction.
Preferably, the burner inner liner is surrounded on the surrounding of the turbine, the outlet end of the burner inner liner and stators and
Turbine rotor communicates, and arc interconnecting piece is equipped at the barrel to outlet end of the burner inner liner.
Preferably, the diameter of the outlet end of the burner inner liner is less than the diameter at the end wall.
Preferably, the turbine is axial type turbine, and the turbine wheel shaft is set to the central axis of the turbine, and described
Turbine drives the compressor rotation by the turbine wheel shaft.
Preferably, the turbine rotor includes blade and wheel disc, and the blade and the wheel disc are integrally formed.
Preferably, the distance between blade of the turbine rotor is gradually reduced by central axis to the wheel disc surrounding.
Preferably, the blade of stators, wheel disc and machine lock are integrally formed.
Preferably, the jet pipe includes the first cylinder successively welded, the second cylinder, third cylinder and the 4th cylinder,
First cylinder is docked with stators machine lock, and rectification supporting plate and whole is provided in first cylinder and second cylinder
Flow cone, first cylinder to the 4th intracorporal flow channel gradually taper up.
Height provided by the invention pushes away the small-size turbojet engine of ratio, compared with prior art, has the following beneficial effects:
1, the distance between blade of compressor is gradually increased by central axis to surrounding, is expanding channel, and air-flow is logical
The deceleration pressurization of expanding channel is crossed, that is, converts kinetic energy into pressure energy, improves air-flow and flow through the pressure of compressor, to increase
The compression ratio of engine；Combustion chamber uses toroidal combustion chamber, and high combustion efficiency, outlet temperature field distribution is uniform, makes stators
It is small with the hot internal stress of being heated evenly for rotor, generation.Fuel oil is sprayed into from burner inner liner rear, rationally utilizes structure, mitigates weight,
Air enters from burner inner liner barrel and fuel oil burns.By the air inlet on arrangement burner inner liner barrel, so that combustion chamber
Maximizing combustion efficiency, reaches highest outlet temperature in the resistance to heat rating that material can be born, and combustor exit temperature height is
Turbine entrance temperature temperature is higher, and the thrust that engine finally generates is also bigger；The compression that above structure passes through raising engine
Increase thrust than the efficiency of combustion with strengthen burning room, to obtain the high effect for pushing away ratio.
2, present invention is specifically directed to small aero fields, can satisfy target drone and carry out the requirement that high maneuver uses,
Engine has the advantages that light-weight, starting is simple, thrust ratio is big simultaneously.
Detailed description of the invention
In order to more clearly explain the embodiment of the invention or the technical proposal in the existing technology, to embodiment or will show below
There is attached drawing needed in technical description to be briefly described, it should be apparent that, the accompanying drawings in the following description is only this
Some embodiments of invention for those of ordinary skill in the art without creative efforts, can be with
It obtains other drawings based on these drawings.
Fig. 1 is the structural schematic diagram of existing combustion chamber flame drum；
Fig. 2 is the structural schematic diagram for the small-size turbojet engine that height of the present invention pushes away ratio；
Fig. 3 is the structural front view of the rotor of compressor；
Fig. 4 is the structure top view of the rotor of compressor；
Fig. 5 is the structural schematic diagram of combustion chamber；
Fig. 6 is the structural schematic diagram of jet pipe of the invention.
1, air intake duct in figure；2, compressor rotor；201, primary blades；202, splitterr vanes；3, turbine wheel shaft；4, turbine turns
Son；5, stators；6, fuel nozzle；7, combustion chamber；8, burner inner liner；81, air inlet；82, curved transition sections；9, radial rectification
Device；10, axial rectifier；11, jet pipe；111, the first cylinder；112, the second cylinder；113, third cylinder；114, the 4th
Body；12, supporting plate is rectified；13, fiaring cone；14, compressor casing；15, fuel injector.
Specific embodiment
To make the object, technical solutions and advantages of the present invention clearer, technical solution of the present invention will be carried out below
Detailed description.Obviously, described embodiments are only a part of the embodiments of the present invention, instead of all the embodiments.Base
Embodiment in the present invention, those of ordinary skill in the art are obtained all without making creative work
Other embodiment belongs to the range that the present invention is protected.
In the description of the present invention, it is to be understood that, term " center ", " length ", " width ", " height ", "upper",
The orientation of instructions such as "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outside", " side " or
Positional relationship is to be based on the orientation or positional relationship shown in the drawings, and is merely for convenience of description of the present invention and simplification of the description, without
It is that the equipment of indication or suggestion meaning or element must have a particular orientation, be constructed and operated in a specific orientation, therefore not
It can be interpreted as limitation of the present invention.In the description of the present invention, unless otherwise indicated, the meaning of " plurality " is two or two
More than.
Its structure of the combustion chamber of existing turbojet engine is as shown in Figure 1, Fig. 1 is the structure of existing combustion chamber flame drum
Schematic diagram；Solid arrow indicates air current flow direction in figure, and dotted arrow indicates fuel oil injection direction；Air-flow air inlet is set to
At end wall, burner inner liner 8 is flowed by 8 end wall of burner inner liner by the air-flow of compressor outflow；Fuel injector is set on barrel,
The injection direction of fuel injector 15 is the radial direction of burner inner liner 8；The drawbacks of above structure, is: the mixing of fuel oil and gas
There are stagnant area, fuel oil is run to the promotion for being largely dependent upon air-flow at turbine, the contact area of air-flow and fuel oil object
It is limited, there is a problem of that burning is insufficient, efficiency of combustion is lower.
The present invention provides a kind of high small-size turbojet engines for pushing away ratio, and referring to Fig. 2-Fig. 5, Fig. 2 is that height of the present invention pushes away ratio
Small-size turbojet engine structural schematic diagram, Fig. 3 is the structural front view of the rotor of compressor, and Fig. 4 is the rotor of compressor
Structure top view, Fig. 5 is the structural schematic diagram of combustion chamber；Including air intake duct 1, compressor, combustion chamber 7, turbine and jet pipe
11, it include spaced primary blades 201 and splitterr vanes 202 on the rotor of compressor, it is adjacent that splitterr vanes 202 are set to two
Primary blades 201 between, primary blades 201 are distorted shape, and the distance between primary blades 201 by the central axis of rotor to surrounding
It is gradually increased (as shown in Figure 3 and Figure 4)；Compressor rotor 2 is connect by turbine wheel shaft 3 with turbine rotor 4；Combustion chamber is annular
Combustion chamber 7 comprising the end wall of burner inner liner 8 and fuel nozzle 6, burner inner liner 8 is opposite with compressor and it is arc, and fuel oil sprays
Mouth 6 is set on end wall, and multiple air inlets 81 are provided on the side wall of burner inner liner 8.
It should be understood that its structure of existing turbojet engine include air intake duct, compressor, combustion chamber, turbine and
Jet pipe, this is several big composition parts of engine, the difference is that internal structure.To specifically connecting for above-mentioned existing structure
The relationship of connecing repeats no more, and if inlet casing front and airplane intake 1 cooperate, rear portion and compressor cooperate；Compressor casing 14
Front is connected with inlet casing, and rear portion is connected with diffuser casing.
Wherein, as shown in Figure 3 and Figure 4, the distance between primary blades are gradually increased by central axis to the surrounding of rotor, i.e.,
Channel between blade is expanding channel, and above structure can increase the pressure of air-flow, i.e. the compression ratio of increase compressor is beneficial
In the raising of motor power.Its cardinal principle are as follows: air intake duct 1 is connected to the middle part of compressor rotor 2, and air-flow is between the blades
Flowing, flow direction are by gradually flowing to surrounding in the middle part of rotor, and since the distance between blade is increasing, air-flow exists
Flow velocity when flowing out rotor is reduced compared to flow velocity when flowing into.Therefore according to Bernoulli equation: p+1/2 ρ v2+ ρ gh=C, formula
Middle p is the pressure of certain point in fluid, and v is the flow velocity of the fluid point, and ρ is fluid density, and g is acceleration of gravity, and h is point place
Highly, C is a constant, it is seen that when fluid flow rate reduces, is increased according to above-mentioned Bernoulli equation air pressure.
It should be noted that calculating in the identical situation of blade dimensions with splitterr vanes and without the leaf of splitterr vanes
Performance curve is taken turns, calculating revolving speed is respectively 20000rpm, 45000rpm, 60000rpm.At the low rotational speed, splitterr vanes are to impeller
Condition range influence less, and at high speed, effective discharge range is essentially identical, have the pressure ratio of splitterr vanes be greater than without point
The pressure ratio of blade is flowed, and the impeller isentropic efficiency with splitterr vanes is apparently higher than no splitterr vanes impeller.And due to of the invention
Its groundwork revolving speed of turbojet engine is in 60000rpm or so, and therefore, setting divergent-centrifugal impeller is conducive to improve this engine
Pressure ratio and working efficiency.
The blade distance of compressor of the invention uses expanding channel, and adds splitterr vanes, and above-mentioned setting effectively mentions
The high compression ratio of compressor, to be conducive to improve the thrust of engine.
Wherein, referring to figure, the fuel oil of above-mentioned combustion chamber 7 is provided by fuel nozzle 6, and fuel oil is axial by the end wall of burner inner liner 8
Injection, air-flow by barrel many places air inlet 81 flow into mixing (in Fig. 5 arrow direction indicate air-flow flow direction), air-flow with
The contact area of fuel oil is big, and 7 efficiency of combustion of combustion chamber of above structure is good, and pitot loss is small, exit flow field and thermo parameters method
Uniformly；And structure is simple, durability is good, light-weight；7 axial dimension of combustion chamber is small, is conducive to the length and drop that reduce turbine wheel shaft 3
Low engine overall weight.
Height provided by the invention pushes away the small-size turbojet engine of ratio, and first, the distance between blade of compressor is in
It is gradually increased at mandrel to surrounding, is expanding channel, air-flow is slowed down by expanding channel to be pressurized, that is, converts kinetic energy into pressure
Power energy improves air-flow and flows through the pressure of compressor, to increase the compression ratio of engine；Second, combustion chamber is using annular combustion
Room 7 is burnt, high combustion efficiency, outlet temperature field distribution is uniform, makes being heated evenly for stators 5 and rotor, answers in the heat of generation
Power is small.Fuel oil is sprayed into from 8 rear of burner inner liner, rationally utilizes structure, mitigates weight, air enters from 8 barrel of burner inner liner and fuel oil
It burns.By the air inlet 81 on arrangement 8 barrel of burner inner liner so that the maximizing combustion efficiency of combustion chamber 7, material can
Reach highest outlet temperature in the resistance to heat rating born, 7 outlet temperature height of combustion chamber, that is, turbine entrance temperature temperature is higher, starts
The thrust that machine finally generates is also bigger；Above structure is by improving the compression ratio of engine and the burning effect of strengthen burning room 7
Rate increases thrust, to obtain the high effect for pushing away ratio.
Present invention is specifically directed to small aero fields, can satisfy target drone and carry out the requirement that high maneuver uses, together
When engine have the advantages that it is light-weight, start that simple, thrust ratio is big.
As optional embodiment, air intake duct 1 is Subsonic inlet, and the inside of air intake duct 1 is by entrance to calming the anger
It is gradually tapered up at machine；And air intake duct 1 is internally provided with rectification supporting plate and fiaring cone.
The air intake duct 1 of above structure can increase intake velocity, can satisfy engine under different working condition, different seas
Charge flow rate requirement under degree of lifting.
As optional embodiment, referring to Fig. 2, compressor further includes radial rectifier 9 and axial rectifier 10, radial direction
Rectifier 9 and axial rectifier 10 are set between compressor rotor 2 and combustion chamber 7, successively by the air-flow of compressor rotor 2
By radial rectifier 9 and axial rectifier 10, directional velocity changes into axial direction.
Air velocity direction by compressor rotor 2 is radial direction, and rectifier 9 radial first is rectified, so that air-flow
No longer disorder, then axial direction is changed into gas velocity direction via axial rectifier 10, in order to enter in next step combustion chamber 7 with
Fuel oil mixing.
The kinetic energy of air-flow is converted pressure energy by compressor: by the gas of compressor can speedup pressurization, speedup be because
To caused by air-flow acting, pressurization is because being formed between the every two adjacent blades of compressor rotor 2 for compressor impeller rotation
The channel of air-flow, and channel is to gradually expand, air-flow, which flows through expansion shape channel, can slow down pressurization, and by kinetic energy transfer pressure
Energy.Preferable scheme is that the channel between radial rectifier 9 and two blades of axial rectifier 10 is also expansion shape, gas
Stream flows through the pressurization that equally can also slow down.
As optional embodiment, referring to fig. 2 and Fig. 5, burner inner liner 8 are surrounded on the surrounding of turbine, the outlet of burner inner liner 8
End is communicated with stators 5 and turbine rotor 4, and arc interconnecting piece is equipped at the barrel to outlet end of burner inner liner 8.
The distribution mode of above-mentioned annular makes 8 outlet temperature field distribution of burner inner liner uniform, make stators 5 and rotor by
Hot uniform, the hot internal stress of generation is small；The arc interconnecting piece in exit can reduce stagnant area, so that the mixing of air-flow and fuel oil
Object more smoothly flows into turbine rotor 4 and stators.
As optional embodiment, referring to fig. 2 and Fig. 5, the diameter of the outlet end of burner inner liner 8 are less than straight at end wall
Diameter.
The bore of outlet end reduces the flow velocity that can effectively improve air-flow Yu fuel oil fluid, reduces the loss of air-flow, so that
The thermal efficiency is higher.
As optional embodiment, turbine is axial type turbine, and turbine wheel shaft 3 is set to the central axis of turbine, and whirlpool
Wheel drives compressor rotation by turbine wheel shaft 3.The process that air-flow flows through turbine rotor 4 can do work to turbine rotor 4, to drive
Turbine rotor 4 rotates, and turbine rotor 4 is connected by turbine wheel shaft 3 and compressor rotor 2, to realize turbine rotor 4 with action pneumatic
Machine rotor 2 rotates.
As optional embodiment, turbine rotor 4 includes blade and wheel disc, and blade and wheel disc are integrally formed.Turbine turns
Son 4 is blade and wheel disc unitary design, and structure is simple, without considering the connection of blade and wheel disc.
As optional embodiment, the distance between the blade of turbine rotor 4 by central axis to wheel disc surrounding gradually
It reduces.
The high temperature and high pressure gas come out from combustion chamber 7 first passes through stators 5 and rectifies to air-flow, makes to flow into rotor
The channel between disorder, the blade of turbine rotor 4 is not gradually reduced by central axis to wheel disc surrounding as gradually airflow direction
Contracting shape channel, air-flow, which flows through, shrinks shape channel meeting speedup decompression.Increasing the main purpose of airflow speed increasing here is to pass through air-flow
It when flowing through turbine rotor 4, does work to turbine rotor 4, flow velocity is easy to that turbine rotor 4 is driven to rotate fastly, and turbine rotor 4 passes through turbine
Axis 3 and compressor rotor 2 connect, to realize that turbine rotor 4 drives compressor rotor 2 to rotate.
As optional embodiment, blade, wheel disc and the machine lock of stators 5 are integrally formed.Likewise, stators
5 blade, wheel disc and casing designs for integral type, and structure is simple, stability is good, without considering the connection of blade and wheel disc.
As optional embodiment, referring to Fig. 6, Fig. 6 is the structural schematic diagram of jet pipe of the invention.Jet pipe 11 wraps
Include the first cylinder 111 successively welded, the second cylinder 112, third cylinder 113 and the 4th cylinder 114, the first cylinder 111 and whirlpool
The docking of 5 machine lock of stator is taken turns, is provided in the first cylinder 111 and the second cylinder 112 and rectifies supporting plate 12 and fiaring cone 13, first
Flow channel in 111 to the 4th cylinder 114 of body gradually tapers up.
Each section of above-mentioned jet pipe 11 can be connected with each other by welding, by 111 to the 4th cylinder 114 of the first cylinder
Flow channel gradually tapers up, i.e., jet pipe 11 is convergent contour jet pipe 11.It can be obtained most suitable by finite element analysis computation
11 discharge area size of convergent contour jet pipe.Above-mentioned flow channel gradually tapers up, and can reduce air-flow in jet pipe 11
Flow losses, to increase the thrust of last outlet end.Wherein, it rectifies supporting plate 12 and fiaring cone 13 is existing device, this field
Technical staff can select as needed, and its structure will not be repeated herein；By reasonably arranging, it may make air-flow in jet pipe
No longer disorder in 11 is conducive to obtain higher thrust.
The height of offer of the invention pushes away the small-size turbojet engine of ratio, its working principle is that: engine is using air as work
Make medium.Air flows into compressor through air intake duct, and compressor blade makes the pressure, temperature and kinetic energy of air-flow to air-flow acting
It changes, and part kinetic energy is further changed by pressure energy by stator blade.The high pressure gas flowed out from compressor exists
In combustion chamber and the fuel oil of fuel system penetrating mixes combining combustion, generates high temperature, high-pressure gas；High-temperature high-pressure fuel gas flows into and pressure
The partial heat energy of turbine of the mechanism of qi on same axis, combustion gas is changed into mechanical energy by turbine, drives compressor rotation；From
The high-temperature high-pressure fuel gas of turbine outflow continues to expand in jet pipe, sprays along engine shaft to from spout high speed outward, makes to send out
Motivation obtains thrust.
In the work above process, lead to first, being improved to expanding by the channel between the blade by compressor rotor
Road, air-flow flows through deceleration pressurization, to increase the compression ratio of engine；Second, being carried out by the structure to combustion chamber flame drum
Design, fuel nozzle are set on end wall, and multiple air inlets are arranged on the barrel of burner inner liner, improve the burning effect of combustion chamber
Rate improves outlet temperature, and combustor exit temperature height, that is, turbine entrance temperature temperature is higher, and the thrust that engine finally generates is also
It is bigger；The combustion chamber of above structure can reduce self weight simultaneously；Third, jet pipe uses multistage shrinkage type barrel soldering, flowing
Channel gradually tapers up, and can reduce flow losses of the air-flow in jet pipe, to increase last thrust.Above structure phase interworking
Close, motor power increase, weight reduce, therefore specific embodiment of the invention part provide turbojet engine can obtain compared with
High thrust ratio.
In the description of this specification, specific features, structure or feature can in any one or more embodiments or
It can be combined in any suitable manner in example.
The above description is merely a specific embodiment, but scope of protection of the present invention is not limited thereto, any
Those familiar with the art in the technical scope disclosed by the present invention, can easily think of the change or the replacement, and should all contain
Lid is within protection scope of the present invention.Therefore, protection scope of the present invention should be based on the protection scope of the described claims.
Claims (10)
1. a kind of high small-size turbojet engine for pushing away ratio, including air intake duct, compressor, combustion chamber, turbine and jet pipe, feature
It is,
It include spaced primary blades and splitterr vanes on the rotor of the compressor, the primary blades are distorted shape, and institute
The distance between primary blades are stated to be gradually increased by central axis to the surrounding of rotor；Compressor rotor is turned by turbine wheel shaft and turbine
Son connection；
The combustion chamber is toroidal combustion chamber comprising burner inner liner and fuel nozzle, the end wall of the burner inner liner are calmed the anger with described
Machine is opposite and it is arc, and the fuel nozzle is set on the end wall, is provided on the side wall of the burner inner liner multiple
Air inlet.
2. the small-size turbojet engine that height according to claim 1 pushes away ratio, which is characterized in that the air intake duct is subsonic speed
The inside of air intake duct, the air intake duct is gradually tapered up by entrance to the compressor；And the air intake duct is internally provided with
Rectify supporting plate and fiaring cone.
3. the small-size turbojet engine that height according to claim 1 pushes away ratio, which is characterized in that the compressor further includes diameter
To rectifier and axial rectifier, the radial direction rectifier and the axial rectifier be set to the compressor rotor with it is described
Between combustion chamber, the radial rectifier and the axial rectifier are successively passed through by the air-flow of the compressor rotor,
Directional velocity changes into axial direction.
4. the small-size turbojet engine that height according to claim 1 pushes away ratio, which is characterized in that the burner inner liner is surrounded on institute
The surrounding of turbine is stated, the outlet end of the burner inner liner is communicated with stators and turbine rotor, and the barrel of the burner inner liner is extremely
Arc interconnecting piece is equipped at outlet end.
5. the small-size turbojet engine that height according to claim 4 pushes away ratio, which is characterized in that the outlet end of the burner inner liner
Diameter be less than the end wall at diameter.
6. the small-size turbojet engine that height according to claim 1 pushes away ratio, which is characterized in that the turbine is axial type whirlpool
Wheel, the turbine wheel shaft is set to the central axis of the turbine, and the turbine drives the compressor by the turbine wheel shaft
Rotation.
7. the small-size turbojet engine that height according to claim 6 pushes away ratio, which is characterized in that the turbine rotor includes leaf
Piece and wheel disc, the blade and the wheel disc are integrally formed.
8. the small-size turbojet engine that height according to claim 7 pushes away ratio, which is characterized in that the blade of the turbine rotor
The distance between be gradually reduced by central axis to the wheel disc surrounding.
9. the small-size turbojet engine that height according to claim 6 pushes away ratio, which is characterized in that blade, the wheel of stators
Disk and machine lock are integrally formed.
10. the small-size turbojet engine that height according to claim 1 pushes away ratio, which is characterized in that the jet pipe include according to
The first cylinder, the second cylinder, third cylinder and the 4th cylinder of secondary welding, first cylinder are docked with stators machine lock,
Rectification supporting plate and fiaring cone, first cylinder to the 4th cylinder are provided in first cylinder and second cylinder
Interior flow channel gradually tapers up.
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| CN201811421442.8A CN109268168A (en) | 2018-11-26 | 2018-11-26 | Height pushes away the small-size turbojet engine of ratio |
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| CN201811421442.8A CN109268168A (en) | 2018-11-26 | 2018-11-26 | Height pushes away the small-size turbojet engine of ratio |
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| CN110159431A (en) * | 2019-05-24 | 2019-08-23 | 长沙市云智航科技有限公司 | A kind of turbo-shaft engine of multi cylinder |
| CN110273782A (en) * | 2019-07-12 | 2019-09-24 | 南昌航空大学 | A kind of Micro Turbine Jet Engine of two-sided NEW TYPE OF COMPOSITE impeller |
| CN111810243A (en) * | 2020-07-17 | 2020-10-23 | 南昌航空大学 | Compressor-turbine integrated engine |
| CN111894680A (en) * | 2020-06-28 | 2020-11-06 | 南昌航空大学 | Miniature turbojet engine with circulating supercharging composite impeller |
| CN112163292A (en) * | 2020-09-24 | 2021-01-01 | 北京航空航天大学 | Ribbed partition nozzle modification method for improving acoustic energy dissipation |
| CN112431686A (en) * | 2020-11-20 | 2021-03-02 | 北京动力机械研究所 | A culvert spray tube for high pressure turbine blade vibration stress measurement tester |
| CN113107679A (en) * | 2021-04-23 | 2021-07-13 | 浙江浙能技术研究院有限公司 | Transition section part for composite tangential air inlet of small gas turbine |
| CN113389759A (en) * | 2021-06-21 | 2021-09-14 | 郭孝国 | Injection device |
| CN113685370A (en) * | 2021-09-24 | 2021-11-23 | 李媛 | High negative pressure steam compressor and impeller |
| CN115095396A (en) * | 2022-06-24 | 2022-09-23 | 西安航天动力研究所 | Turbine outlet diversion elbow structure of liquid rocket engine |
| CN115213844A (en) * | 2020-07-02 | 2022-10-21 | 中国航发常州兰翔机械有限责任公司 | Centrifugal impeller assembly of aviation turbine starter and assembling method thereof |
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| CN111894680A (en) * | 2020-06-28 | 2020-11-06 | 南昌航空大学 | Miniature turbojet engine with circulating supercharging composite impeller |
| CN115213844A (en) * | 2020-07-02 | 2022-10-21 | 中国航发常州兰翔机械有限责任公司 | Centrifugal impeller assembly of aviation turbine starter and assembling method thereof |
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| CN113107679A (en) * | 2021-04-23 | 2021-07-13 | 浙江浙能技术研究院有限公司 | Transition section part for composite tangential air inlet of small gas turbine |
| CN113107679B (en) * | 2021-04-23 | 2023-09-19 | 浙江浙能技术研究院有限公司 | Transition section component of composite tangential air inlet of small-sized gas turbine |
| CN113389759A (en) * | 2021-06-21 | 2021-09-14 | 郭孝国 | Injection device |
| CN113685370A (en) * | 2021-09-24 | 2021-11-23 | 李媛 | High negative pressure steam compressor and impeller |
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