This invention relates to turbine engines, and to a method of preventing the turbine of a turbine engine from choking at high speed. More particularly, it relates to novel improvements to prevent air from compressor discharge to bypass the combustion system. The present invention concerns gas turbine engines for auxiliary power units on aircraft, spacecraft, missiles, and other vehicles.
A typical turbine scroll system is shown in
Current needs for turbine scroll systems include the ability to control small amounts of gas leakage between components at various operating conditions for performance optimization. Two main operating conditions are an open-loop condition (e.g., ground maintenance or in-flight emergency power) in which the engine runs on its own power and a closed-loop condition (e.g., taxi condition and general flight conditions) in which the engine runs on the bleed gas of the main engine.
Being able to control the size of the B-width (gap between the combustor scroll 240 and associated structures to minimize the gas leakage that can adversely effect the engine performance) may be a concern during engine design and development. Controlling the size of the B-width becomes more critical when an engine is operated in dual modes. Keeping the size of the B-width constant and maintaining effective sealing for system performance and integrity is critical during all operating conditions and surges. A constant B-width size minimizes gas leakage and erosion between components that may cause excessive wear or fretting. Prior art systems usually require tight tolerances for the inner diameters and outer diameters of mating components and shields around surfaces to minimize leakage.
None of the prior art is specifically intended for high performance, high cycle applications, and some suffer from one or more of the following disadvantages:
a) excessive wear and fretting.
b) inability to maintain constant B-width size.
c) ineffective sealing.
d) gas leakage between components.
As can be seen, there is a need for an improved apparatus and method for an improved gas turbine engine system, which minimizes wear and fretting, maintains constant B-width size, and effectively seals to prevent gas leakage between components.
In one aspect of the present invention, a gas turbine engine comprises: a turbine scroll inside a combustor housing; a forward discourager; an aft discourager; a B-width situated between the forward discourager and the aft discourager; a forward bayonet situated on the forward side of the turbine scroll; a radial nozzle contacting the forward bayonet on the forward side of the turbine scroll at a bayonet engagement point; an aft scroll ring; a retaining ring securing the turbine scroll while maintaining an axial loading point on the aft scroll ring; a forward scroll ring and the retaining ring restraining displacement of the forward scroll ring and the aft scroll ring.
In another aspect of the present invention, a gas turbine engine comprises: a turbine scroll inside a combustor housing; a forward discourager; an aft discourager; a B-width situated between the forward discourager and the aft discourager; a forward axial seal adjacent to the forward discourager; an aft axial seal adjacent to the aft discourager; the forward discourager comprising a 90 degree bending angle; the aft discourager comprising a 90 degree bending angle; a radial nozzle engaged with a forward bayonet on the forward side of the turbine scroll; the forward bayonet contacting the radial nozzle at a bayonet engagement point; an aft scroll ring; a retaining ring adjacent the aft scroll ring; the retaining ring securing the turbine scroll while maintaining an axial loading point on the aft scroll ring; a forward scroll ring; and the retaining ring restraining displacement of forward scroll ring and the aft scroll ring.
In a further aspect of the present invention, a gas turbine engine comprises: a turbine scroll inside a combustor housing; the turbine scroll comprising four pairs of sealing surfaces; a B-width situated between the forward discourager and the aft discourager; a forward bayonet adjacent the forward side of the turbine scroll; the forward bayonet contacting the radial nozzle at a bayonet engagement point a retaining ring adjacent an aft scroll ring; the retaining ring securing the turbine scroll while maintaining an axial loading point on the aft scroll ring; a forward scroll ring; and the retaining ring restraining displacement of the forward scroll ring and the aft scroll ring.
In yet another aspect of the present invention, a gas turbine engine comprises: a compressor section; a combustor section; a compressor scroll; a turbine scroll inside a combustor housing; the turbine scroll comprising four pairs of sealing surfaces; a forward discourager; an aft discourager; a B-width situated between the forward discourager and the aft discourager; a forward axial seal adjacent to the forward discourager; an aft axial seal adjacent to the aft discourager; the forward discourager and the aft discourager comprising a 90 degree bending angle for flow restriction; a radial nozzle engaged with a forward bayonet on the forward side of the turbine scroll in six locations; the forward bayonet contacting the radial nozzle at a bayonet engagement point; a radial seal on the forward side of the B-width; a radial seal on the aft side of the B-width; a retaining ring adjacent an aft scroll ring; the retaining ring securing the turbine scroll while maintaining an axial loading point on the aft scroll ring; a forward scroll ring; and the retaining ring restraining displacement of forward scroll ring and the aft scroll ring.
In another aspect of the present invention, a method is disclosed for preventing a gas turbine engine of an auxiliary power unit from choking at high speed comprises: introducing a portion of the exhaust gas of an associated turbine engine through a radial nozzle; maintaining a constant B-width; securing a retaining ring on the aft side of a turbine scroll while maintaining an axial loading point on the aft side of a scroll ring; and restraining displacement of the turbine scroll by the retaining ring.
These and other aspects, objects, features and advantages of the present invention, are specifically set forth in, or will become apparent from, the following detailed description of a preferred embodiment of the invention when read in conjunction with the accompanying drawings.
The following detailed description is of the best currently contemplated modes of carrying out the invention. The description is not to be taken in a limiting sense, but is made merely for the purpose of illustrating the general principles of the invention, since the scope of the invention is best defined by the appended claims.
The present invention is useful for auxiliary power units for all types of flight vehicles, including, but not limited to, aircraft, missiles, and spacecraft. As opposed to the prior art gas turbine engine shown in
An exemplary gas turbine engine 200 according to the present invention is shown in
As in
Referring to
With reference to
Referring back to
The gas flow leakage D on the four pairs of sealing surfaces (320, 330; 340, 350; 360, 370; 380, 390) seals both the forward and aft sides 300, 310 of the turbine scroll 10 while maintaining constant the B-width 110 of the scroll 10. Axial seal 60, 90 surfaces, shown in
A method for preventing a gas turbine engine of an auxiliary power unit from choking at high speed may comprise diverting a portion of the exhaust gas of an associated turbine engine by about 90 degrees in flow direction; maintaining a constant B-width 110; introducing the diverted exhaust gas through a radial nozzle 30; securing a retaining ring 50 on the aft side 310 of a turbine scroll 10 while maintaining an axial loading point 130 on the aft side 310 of a scroll ring by direct pressure; restraining displacement of the turbine scroll 10 by the retaining ring 50; and, further reinforcing contact at an axial loading point 130 on the aft side 310 of the scroll ring by direct pressure.
Although the present invention has been described in considerable detail with reference to certain preferred versions thereof, other versions are possible. Therefore, the spirit and scope of the appended claims should not be limited to the description of the preferred versions contained therein.
The invention was made with Government support under contract number N00019-01-C-3002 with outside funding from Lockheed Martin—US Government under Joint Strike Fighter (JSF) Program. The government has certain rights in this invention.
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Number | Date | Country | |
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20050076643 A1 | Apr 2005 | US |