WO1996013656A1

WO1996013656A1 - Gas turbine engine

Info

Publication number: WO1996013656A1
Application number: PCT/BG1995/000011
Authority: WO
Inventors: Todor Georgiev Givetchev
Original assignee: Todor Georgiev Givetchev
Priority date: 1994-11-01
Filing date: 1995-10-31
Publication date: 1996-05-09
Also published as: BG99154A

Abstract

A gas turbine of symmetrical configuration comprises on the ends of a shaft (1) two bearings (2), two axial compressors (5), two recuperation chambers (10) and centrally two radial compressors (11) arranged back to back. Air enters from the air inlets situated at the two ends of the engine and passes through the axial compressors (5) via the recuperation chambers (11) and the radial compressors (5) to central combustion chambers (14). The exhaust gases then flow from the central combustion chambers via two axial turbines (18) to two exhaust ducts. The turbines (18) are concentric to the compressors (5) and are situated on the outside of the recuperation chambers (11). Cooling air can be fed via holes (19) from the recuperation chambers (11) to the turbine blades (8).

Description

GAS TURBINE ENGINE

This invention relates to gas turbine engine, applicable in car building indastry.in aviation indastry and especially in helicopter production.

Gas turbine engine of common type is known.comprising two shafts.both installed on one and the same axle, rotating in opposite directions. At the front end of the first shaft a low pressure comressor is mounted and then followed by high pressure compressor. Low pressure compressor blades are situated within the fuselage as to split the gas flow comming out of the low pressure compressor into two streams.The first one passing through high pressure compressor and the second sream passing above its casing . Inbetween the high pressure comressor and the turbine which is mounted at the end of the first shaft and combustion chamber is installed.At the front end of the second shaft a second turbine is mounted whose upward part is blown by the second stream comming out of low pressure compressor and its downward part is blown by the first stream comming out of a turbine mounted at the end of the first shaft. The construction is complex.heavy .massive with complex bearing and and heavy use of cealings.As a result of unsufficient refregerating of the turbine blades they are worn out quickly, which leads to a frequent dismantle. Disclosure of invention Gas turbine engine comprises a shaft.bearings at its both ends installed in fulcrums .followed by symmetrically mounted left and right axle blade compressors with turbines in a way, that the back sides of the last row of compressor blades are firmly fixed to the bases of the last row of turbine blades. At a distance from the bases of the turbine blades , upwards to their tops, radially to the shaft ,the casing is fixed .which casing devides the turbine blades , respectively to a lower (passive) part and an upper(actιve) part.called "working blades". On the casing openigs are perforated at the contact zones of the bases and the tops of the turbine blades, to the casing itself .excluding the last row of the blades which holes serve the purpose to cool off the working blades of the turbine. The front end of the casing is steadily fixed to the periphery of the last row of the compressor blades. The space inbetween the shaft and the casing, comprising the blade bases, represents respectively the right and left regeneration chambers .Whereafter rotary compressor is mounted and above it combustion chamber is installed devided into two symmetrical parts. In the combustion chamber, burners are installed and right below them the inlet guide vanes are mounted in socket joints .which help the regulation, dosage and cutting off the fuel supply ,and the combustion of the gas coming from the rotary compressor to the left or right side or both sides of the combustion chamber.On the both sides of the combustion chamber the nozzle blades are installed .fixed steadily to the fuselage of the gas turbine engine, which nozzle blades guide the gas flow. Right after the nozzle blades.the turbine blades and the inlet guide vanes are installed. The said inled guide vanes are fixed to the outer casing. After the turbines, outlet ducts are chambered for the worked out gas. Gas turbine engine is of a very light and simple type construction ,not massive and bulky and is of a high prossesing reliability. The high temperature in the combustion chamber (up to 2000 C) ensures maximum expansion of the gas and does noi cause a turbine blade destruction, because the main gas flow, supercharged by the axle compressors , passes trough their bases and assures their maximum cooling off.That is why there is no need for aditional .deliberate refregerating. These improved working conditions ensure long lasting life to the engine. The construction of the gas turbine engine assures full usage of rotary compressor^Ms energy and the energy caused by the expansion of the gas to the both sides of the combustion chamber in the course of the burning process. High working capacity is assured up to 65% at temperature of the surrounding 15 C (228 k). Bearings are mounted at the both ends of turbine shaft, which assures normal processing conditions and quite a low temperature for the bearings.Quite a great number of cealings, typical for the cunstructions known, are avoided.

Implementation of the invention is shown on fig. 1, representing longitudinal section of the gas turbine engine.

Gas turbine engine comprises shaft 1 supplied with bearings 2 installed in dungeon pillows 3 at both ends of the shaft 1 whereafter axial blade compressors 5 are mounted as their inlet guide vanes 4 are fixed steadily to the outer casing 6 . Directly after the compressors 5 the bases of the blades 7 are fixed to the shaft 1. Radially to the shaft 1 and at a distance from the bases of the blades 7,the casing 9 is fixed and to the said casing 9 the tops of the blades 8 of the turbines 18 are fixed.The front end of the casing 9 is fixed to the periphery of the last compressor blade row 5 of the casing and at the zones of contact to the shaft 1 of the bases of the blades 7 .and the working /active/ blades 8, excluding the last row of blades, openings 19 are chambered The space inbetween the shaft 1 and the casing 9 represents the regeneration chamber 10 and right after is the inlet of the rotary compressor 11. Combustion chamber 14 is situated inbetween the rotary conmpressor 11 and the fuselage 17 compresing the burners 13 and the guide vanes \2 are mounted inside. After the combustion chamber 14, symmetrically on the both sides of the fuselage 17 nozzle blades 15 and the guiding vanes 16 of the turbines 18 are mounted. After the turbines 18 the outlet ducts for the worked out gas are mounted.

Gas turbine engine in accordance with the invention works as follows.The air flow comes simultaneously to the inlets of the left and right axial compressors 5 and the inlet guide vanes 4.The air is supercharged and is passed to the regeneration chambers 10 where part of the air flow goes through the chambered holes 19 of the casing 9 and cools off the turbine blades 8 while the main part of the air stream goes through the regeneration chambers 10 and increases its temperature up to 900 C.Through the outles of the regeneration chambers 10 the air flow is drawn in by the rotary compressor 11 where the pressure in increased as well as the cynetic energy of the gas.Passing through the rotary compressor 11 the gas is distributed and regulated by means of the guide vanes 12 to the both sides of the combustion chamber 14. The fuel supplied by the burners 13 due to the increased temperature and pressure of the gas in the combustion chamber 14, burns out and expands the gas, that goes simultaneously to the left and right nozzle device 15 and then to the active turbine blades 8 of the left and right turbines 18 and the turbine guide vanes 16. Transmitting its energy to the turbines 18 the gas flows out in the atmosphere.

Claims

I.Gas turbine engine comprising corpus in which casing, combustion chamber with burners, nozzle ducts and a shaft with bearings are mounted.on the shaft axial compressors are mounted with turbine guide vanes.high pressure compressor and turbines, speciality characterising the shaft IM are the bearings 121 mounted on its both ends where after next are mounted the left and right axial blade compressor /5/ installed to the back sides of the blades of the last row of turbine blades /18/ and at a distasnce, upwards from from the bases of the turbine blades /18/ and radially to the shaft IM the casing is fixed so that the space inbetween the shaft IM and the casing /9/ comprissing the bases of the blades 111 represents the regeneration chamber /10/ whereafter a centrifucal high pressure compressor is mounted/11/ followed by the combustion chamber /14/, formed by the outer casing /17/ part of the inner casing /9/ and the sentrifucal compressor /11/ .comrissing the guide vanes /12/ and the burners/13/ and at the end of the outlets of the combustion chamber /14/ a variable nozzle device /15' is installed, directing the gas flow to the working blades /8/ of the turbines ,'18

2.Gas turbine engine is claimed in claim 1 wherein the casing /9 is perforated at the zones of the contact of the bases of the blaαes 7/ and working blades /8/ excluding the last row of the turbine blades /18/ and the front end of the casing/9/ is fixed to the periphery of the last row of the compressor blades 15/.

3.Gas turbine engine is claimed in claim 1 wherein the guide vanes /12/ which regulates the fue supply and the combustion of the gas coming from the centrifugal compressor to the left or rihgt or to the both sides of the combustion chamber /14/ is installed in a socket joint.