GB2587790A

GB2587790A - Vartle ©

Info

Publication number: GB2587790A
Application number: GB1911594.8A
Authority: GB
Inventors: Mark Edmund Porter Andrew
Original assignee: Individual
Current assignee: Individual
Priority date: 2019-08-13
Filing date: 2019-08-13
Publication date: 2021-04-14
Also published as: GB201911594D0

Abstract

A gas turbine engine in which the casing of the axial compressor section has a porous section or a number of holes 9. These holes may lead to a manifold via air passages, the manifold may consist of a vacuum channel 10 and a return flow channel 11, the vacuum channel may have a non-return valve or reed valve 12 at a number of stages such that the radial air leakage or air spill at each blade tip is recovered and fed into the engine combustion system from a return manifold at higher pressure. The air entering the return channel is of high pressure and is bled off from the compressor. The manifolds may be radially located around the radius of the compressor casing, or may take the form of a single cannular manifold. The whole system is known as VARTLE, this is the acronym for Vacuum Accentuated Radial Turbulence Loss Enhancer.

Description

Reference: Strateghist 1 Title A gas turbine engine with the VARTLE© modification on the casing of the compressor section.

Description of invention

This invention relates to a gas turbine power plant which may be applied to a unit which is either designed specifically for aircraft propulsion or for a land-based gas turbine power plant for electric power generation or for a marine application using similar core components but not necessarily so. By core components we mean 1.A compressor specifically of the axial type of any construction but may include those with a contra-rotating configuration, 2.The total combustion system also known as the combuster and 3. The turbine component used specifically to drive the compressor component which may also include for additional separate turbine stages for power drives via a gearbox or similar. The marine application may include for a direct drive to a propeller propulsion unit or via a generator to provide power for an electric driven propeller propulsion unit.

The dassical pressure/temp cycle for a gas turbine has three basic parameters which determine its power output either as a turbo-jet (with or without a by-pass fan) or as unit for providing pure rotational power. For the purposes of this description they are to be considered the same. These are 1.Mass air flow per unit time, 2 Compression ratio, that is the total pressure rise across the compressor, and 3 Temperature of combustion which determines the energy content or enthalpy available to do useful work by expansion through the turbine or a similar expansion component The actual power output depends on the maximum temperature which is a metallurgical solution, the efficiency of the compressor and turbine and is dependent on the shape of the respective blades and are configured to give a generally theoretical vortex flow which has the advantage of reducing the air or radial leakage past the tip of the blades. The type of compressor under consideration is specifically of the axial type but also included are those which are a combination of both the centrifugal and axial type configured into a single unit which are generally confined to the smaller gas turbine units.. A compressor in this context is specifically one of the axial variety in which the moving blades, which may be made integral with a disc or separately, mounted radially on a rotatable separate set of discs or a drum type rotor.. Each moving blade is associated with a fixed stationary blade integral with the outer compressor casing but not necessarily so. The static blade may or may not be rotatable about its longitudinal axis. A compressor stage is generally defined as comprising a fixed static

Description cont:-

blade located by some means to the outer compressor casing and a moving blade fixed by some means to a rotor shaft driven from a turbine which may be single or multi stage... To obtain the required compression ratio an unspecified number of compressor stages are located on a common rotor or shaft but in spite of the twisted shape to provide the ideal vortex flow there is still an unspecified radial leakage of air at the tips of each blade, or air spill causing a loss of efficiency due to the tip leakage of the air at each stage of the compressor configured into an unspecified number of stages.

According to the present invention there is provided a gas turbine machine consisting of three separate components namely the compressor the combuster (combustion system) and the turbine unit. The Vanle' component is specific to the compressor component and consists of a radially a disposed porous unit or array of holes generally aligned with the axis of rotation of the compressor stages but not necessarily so. An axial compressor stage consists of both a fixed and moving blade. . The number of stages in a compressor is determined by the design compression ratio and the \raffle component can be located at any number of locations along the line of the compressor stages. Located around the porous unit or array of holes is a radially located manifold.

The number of radially disposed manifolds can be any number to a complete annular (or cannular) arrangement surrounding the periphery of the compressor casing from the inlet stage to the final stage. Each complete manifold pipe consists of both a vacuum section commencing at the inlet compressor stage which lies along the entire length of the compressor casing (but not necessarily so) and a superimposed a return flow manifold which carries both the vacuum air the return air. The initial air carried in the vacuum air manifold is defined as spill air which is the leakage of air across the tips of the moving compressor blades. The vacuum air manifold is fitted with a series of typically non-return or reed type valves located along the vacuum pipe to prevent blow back as the pressures increases. At the extremity of the vacuum channel the air is mixed with incoming at a higher pressure from a final or late compressor stage and the combined flow is reversed back into the vacuum channel which then combines with the final stage air before entering the combuster section.

Description cont:-

Statement of invention to accompany drawings

A specific aspect of the invention is that it can be applied to all gas turbine machines of the axial or of the combined centrifugal-axial compressor type whether they are for aircraft propulsion, land or sea based for power generation or for gas compression or for direct marine gas turbine marine electric propulsion..

This particular embodiment will now be described by way of an example.

Figure 1 shows a cross-section representation of a typical gas turbine core unit, comprising an axial compressor 1. combuster 2,and power tmbine 3, mounted typically on a common single spool shaft 4 but not exclusively so..Air 5 at a low pressure enters the compressor intake and is compressed in a number of stages 6. On the inside surface of the compressor casing 7 is located a porous membrane or an array of holes of an unspecified number 8.which are integral with the vacuum channel as a part of the manifold unit 12 the Wale' unit as shown in detail see Figure 2.

In Figure 2 The porous membrane 9 forming part of the inner surface of the compressor housing which can also be an array of holes. The number of holes positioned radially is not specified as it will depend on the size of the compressor and design pressure ratio. The manifold has two air channels. The vacuum channel 10 and the return flow channel 11 Air entering the return channel 14 is of high pressure and is bled off from the compressor discharge air by some means, the pressure of which can be controlled by a pressure reducing non-return valve 16. Pressure in the vacuum manifold 10 is such that the air pressure 14 from the return flow channel 11 is sufficient to force spill air from the radial compressor stage leakage to mix with the return air flow 15 before entering the combustion chamber. Various non return valves or reed type valves 12 are indicated.

The invention described above can be configured as comprising the total Vara? system

Claims

CLAIMS1. A gas turbine compressor of the axial type but not limited to such that has a porous membrane or a number of radial holes configured as a part of a turbine casing.
2. A gas turbine compressor according to claim 1 that has a series of manifolds affixed by some means either integral to or as part of the outer compressor casing.
3. A gas turbine compressor according to claim 2 has a manifold consisting of a vacuum channel and a return flow channel.
4. A gas turbine compressor according to claim 3 has a number of non-return or reed type valves in the vacuum channel which collects the radial air leakage or air spill from each or a number of compressor stages.
S. A gas turbine compressor according to claim 4 has a reverse flow channel with a non-return pressure control valve.
6. A gas turbine compressor according to claim 5 has a number of manifolds radially located on the circumference of the compressor casing but the can also be represented by a single cannular type manifold embracing the circumference and full length of the compressor.
7. A gas turbine compressor according to claim 6 has a reverse flow channel as given in claims 3 and 5 at higher pressure and drives the air along the vacuum channel
8. A gas turbine compressor according to claim 7 that can discharge air into any compressor stage where the total manifold discharge pressure in the vacuum is just greater than the particular compression stage pressure
9. A gas turbine compressor of the axial type but not limited to such covered by claims 1 to 8