CN1225414A - High performance turbine engine - Google Patents
High performance turbine engine Download PDFInfo
- Publication number
- CN1225414A CN1225414A CN 98105004 CN98105004A CN1225414A CN 1225414 A CN1225414 A CN 1225414A CN 98105004 CN98105004 CN 98105004 CN 98105004 A CN98105004 A CN 98105004A CN 1225414 A CN1225414 A CN 1225414A
- Authority
- CN
- China
- Prior art keywords
- turbine
- blade
- compressor
- spacer ring
- turbine engine
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Images
Landscapes
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The present invention relates to a high-performance turbine engine and gas turbine engine which possess the following characteristics: (1). Its special structure not only can fully utilize the fuel gas energy in the turbine, but also strengthens the cooling effect of high temp. parts, enhances integral efficiency, elongates service life. (2). The length of rotor shaft is shortened by 40%, the dynamic balance is raised apparently. (3). The structure of whole machine is rational and compact, cooling system, lubricating system and setting and adjusting procedure are simplified, the production cost is lowered.
Description
The present invention relates to gas turbine engine.Gas turbine engine be a kind of with fuel conversion of heat into kinetic energy and externally continue the device of acting, be widely used in fields such as aviation, navigation, generating and outputting power, it is main the argumentation that this specification starts with the turbine of aviation field.
Aero-turbine is divided into turbojet engine, turbofan engine, turboaxle motor, turbine propelling screw engine, is called for short whirlpool spray, turbofan, whirlpool axle, whirlpool slurry motor.The present invention relates generally to whirlpool axle, whirlpool slurry motor.Problem that aero-turbine is existing and development trend: improve the reliability of overall work, reduce oil consumption and blowdown, prolong the average time between overhauls of each component, optimize structure, thereby further reduce manufacture cost and user cost.
For the functional reliability that improves motor and prolong its average time between overhauls, should reduce the vibration values of motor as far as possible, because vibration is greatly the major reason that causes fatigue ruption, vibrative factor is a lot, wherein mainly be caused by rotor unbalance, because the rotating speed when motor real work rotating speed is much higher than blance test, and during working rotor, temperature height, stress are big, its generation elasto plastic deformation, part are misplaced mutually, dynamic unbalance significantly increases, and is the important measures that reduce vibration so reduce rotor unbalance degree.Must be short with the engine rotor system design, rigidity is bigger, with the distortion of avoiding under thrust, gravity and gyrostatic moment effect, producing, reduce rotor Roughness (foregoing is consulted " aerial turbo fan engine " the 113rd, 119,122,437,448 page that National Defense Industry Press publishes in December, 1985).
Based on above-mentioned analysis, as shown in Figure 1, the structural design that the present invention changes conventional engines goes out particular structure, along on the rotor axial, save the firing chamber and turbine occupies, make axial length shorten 40%, significantly improve rotor rigidity and dynamic balance performance, reduce vibration, thereby improve the functional reliability of motor, and prolong its average time between overhauls.
For high-pressure turbine efficient is improved and weight reduction, adopt preceding stagnation temperature of higher turbine and transonic turbine, the stagnation temperature key adopts advanced cooling technology and high temperature material before improving turbine, because of the development and application of high temperature material limited, so solve the cooling problem of turbine part outstanding (above content is consulted " aerial turbo fan engine " the 157th, 158,427,448 page that National Defense Industry Press publishes in December, 1985).
As shown in Figure 1: the present invention with turbine blade and compressor blade in conjunction with as a whole, the air-flow of cooling turbine can directly be drawn from compressor blade, and can utilize the positive and negative blade face pressure difference of compressor blade, cooling blast is drawn the back pressure mechanism of qi, form the circulation of cooling blast, to reduce the compressor efficiency loss, can strengthen cooling gas flow on this basis, to strengthen cooling effect, both prolonged the working life of high-temperature component, reduce the enthalpy of turbine outlet air-flow again; On the other hand, the present invention is nested in turbine active chamber and gas compressor active chamber in the rotating shaft, and the air in the gas compressor can fully absorb the heat energy of turbine part.Under the constant condition of rotor speed and compressor inlet gas flow temperature, gas flow is constant in the gas compressor, according to: P
1V
1/ T
1=P
2V
2/ T
2, air is compressed in gas compressor in the process, absorbs heat again, to further improve outlet temperature and the pressure ratio of air-flow at gas compressor, thereby improve the total enthalpy of air-flow before turbine, the effect of this two aspect will increase the enthalpy drop value of gas flow in turbine, improve overall efficiency, reduce oil consumption.
Its specific structure of the present invention is compressed in the process air in gas compressor, both absorbed kinetic energy, has absorbed heat energy again, and the existing turbogenerator of the latter can't be realized, be the net added value that improves the complete machine overall efficiency.
The existing cooling blast that starts is to draw from gas compressor by pipeline, walks around the firing chamber, flows to turbine, can't reflux cycle, and not only cause compressor efficiency to descend, and also system complex, easy break-down.In addition, because the rotor axial size is long,, adopt the assembly technology of multistep balance for guaranteeing the functional reliability of motor, cycle is long, the manufacture cost height, the present invention's compactness rational in infrastructure has been simplified cooling system and lubrication system, after rotor length shortens, its rigidity and dynamic balancing degree significantly improve, thereby have simplified the technology of multistep balance, reduce manufacture cost.
In sum, the present invention and existing like product relatively have following characteristics:
1. its its specific structure had both made full use of the combustion gas energy, had strengthened the cooling effect of high-temperature component again, thereby improved overall efficiency, increased the service life.
2. rotor axial length shortens 40%, and its rigidity and dynamic balance performance significantly improve, and not only improve the functional reliability of complete machine, and prolong the average time between overhauls of each part.
3. the said device features rational structure compactness has been simplified cooling system, and lubrication system and assembling and setting technology have reduced manufacture cost.
Accompanying drawing 1 is the symmetrical half sectional view of structure of the present invention.
Accompanying drawing 2 is the air seal structure partial sectional view between turbine active chamber of the present invention and the gas compressor active chamber.
Now be further described by reference to the accompanying drawings:
As shown in Figure 1, turbine moving vane (3) passes through middle with compressor work blade (2) Spacer ring (4) combines, and consists of an integral body, turbine by wheel disc (1) and axle and bearing again Guide vane (5) combines by middle spacer ring (6) with compressor guide vane (7), and Be loaded on the outer casing (7), the annular chamber that outer casing (7) and rotating shaft (10) consist of is by the centre Casing (9), moving vane spacer ring (4) and guide vane spacer ring (6) are divided into inside and outside two jointly Individual concentric annular chamber consists of respectively turbine working chamber and compressor working chamber, and the present invention adopts With baffling formula toroidal combustion chamber.
The present invention can specifically implement in application by the following method.
1. in spacer ring, the guide vane between the spacer ring, all need adopt corresponding air seal structure, as shown in Figure 2 in middle casing and the moving vane according to airflow direction.
2. for reducing manufacture cost, turbine blade can be selected different materials for use with compressor blade.
3. for reducing the motor radial dimension, can adopt the return flow type annular combustion chamber.
4. effectively work in the speed range of broad for the assurance motor, compressor inlet place guide vane can be made the guide vane of angle adjustable.
5. do corresponding change as with the exchange of the turbine active chamber shown in the accompanying drawing 1 and gas compressor active chamber position, and with compressor inlet, position, firing chamber and blade shape, can develop into the core engine of turbofan engine.
The data that the present invention consults:
1. National Defense Industry Press's in December, 1985 publication " aerial turbo fan engine) "
2. National Defense Industry Press publishes [English] Luo Ersi Roy Si company in October, 1975 and compiles " air breathing engine "
3. aviation industry publishing house publishes " World Airways Engine Manual " in June, 1996
Claims (3)
- The present invention relates to gas turbine engine, particularly motor is starched in whirlpool axle, whirlpool, it is characterized in that:1. turbine moving blade and compressor blade and blade radially are combined into an integral body by middle spacer ring, link by wheel disc and rotating shaft again, and during work, the kinetic energy that turbine moving blade absorbs is directly delivered to compressor blade and blade.
- 2. turborotor and gas compressor guide vane radially are combined into an integral body by middle spacer ring, are loaded on outer casing.
- 3. outer casing is divided into inside and outside two concentric annular chambers with the annular chamber that rotating shaft is constituted jointly by the middle spacer ring of the middle spacer ring of middle casing, moving vane and guide vane, constitutes turbine active chamber and gas compressor active chamber respectively.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN 98105004 CN1225414A (en) | 1998-02-02 | 1998-02-02 | High performance turbine engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN 98105004 CN1225414A (en) | 1998-02-02 | 1998-02-02 | High performance turbine engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN1225414A true CN1225414A (en) | 1999-08-11 |
Family
ID=5218631
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN 98105004 Pending CN1225414A (en) | 1998-02-02 | 1998-02-02 | High performance turbine engine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN1225414A (en) |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101981276A (en) * | 2008-03-25 | 2011-02-23 | 涡轮梅坎公司 | Turbine engine including a reversible electric machine |
| CN106677902A (en) * | 2015-11-05 | 2017-05-17 | 熵零股份有限公司 | Turbocharger |
| CN108825380A (en) * | 2018-05-28 | 2018-11-16 | 华中科技大学 | A kind of high efficiency turboshaft engine |
| CN110005644A (en) * | 2018-01-04 | 2019-07-12 | 中国航发商用航空发动机有限责任公司 | Axial flow compressor stator with intermediate casing |
-
1998
- 1998-02-02 CN CN 98105004 patent/CN1225414A/en active Pending
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101981276A (en) * | 2008-03-25 | 2011-02-23 | 涡轮梅坎公司 | Turbine engine including a reversible electric machine |
| CN101981276B (en) * | 2008-03-25 | 2014-09-10 | 涡轮梅坎公司 | Turbine engine including a reversible electric machine |
| CN106677902A (en) * | 2015-11-05 | 2017-05-17 | 熵零股份有限公司 | Turbocharger |
| CN110005644A (en) * | 2018-01-04 | 2019-07-12 | 中国航发商用航空发动机有限责任公司 | Axial flow compressor stator with intermediate casing |
| CN108825380A (en) * | 2018-05-28 | 2018-11-16 | 华中科技大学 | A kind of high efficiency turboshaft engine |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN104675510B (en) | A kind of quick response high-altitude two-stage turbocharger of low pneumatic inertia | |
| CN104196571B (en) | A kind of method and its device for improving turbine engine efficiency | |
| CN101021179A (en) | Turbosupercharger axle radial flow air compressor structure | |
| CN206468433U (en) | A kind of engine mechanical supercharger | |
| CN101078354A (en) | Porous metal vane coupling design method | |
| KR20140014175A (en) | A turbine wheel, a turbine and use thereof | |
| CN104963775A (en) | Miniature type bladeless gas turbine | |
| CN100494692C (en) | Compact pneumatic overall arrangement of compression system in aerial turbo fan engine | |
| CN1225414A (en) | High performance turbine engine | |
| CN110360157B (en) | Gas turbine compressor cascade shroud of spatial structure | |
| CN1693692A (en) | Miniature single propeller turbine fan engine | |
| CN2147357Y (en) | J80 turbo supercharge and application thereof | |
| He et al. | Experimental investigation of the sealing performance of honeycomb seals | |
| CN110821570B (en) | Gas turbine without main shaft | |
| JP4448854B2 (en) | Turbo compressor system for an internal combustion engine with an impeller comprising a radial type compressor and having a swept wing | |
| CN210152737U (en) | Volute component of exhaust gas turbocharger | |
| CN201982168U (en) | Dynamic self-regulation turbocharger | |
| CN212272608U (en) | Turbine type air compressor suitable for fuel cell system | |
| CN202132103U (en) | Turbocharger | |
| CN101469635A (en) | Turbocharger | |
| CN112096509A (en) | Streamline tunnel type turbocharger adopting outline type air bearing | |
| CN111472879A (en) | Novel two-stage supercharger | |
| CN112780400B (en) | Special booster for high-altitude power recovery of aviation piston engine | |
| CN1031422C (en) | Gas-wave pressurizer | |
| CN215109565U (en) | Waste gas energy recovery type air compressor for fuel cell |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| C01 | Deemed withdrawal of patent application (patent law 1993) | ||
| WD01 | Invention patent application deemed withdrawn after publication |