Patent Description:
Gas turbine engines, such as those which power modern commercial and military aircraft, include a compressor section, combustor section and turbine section arranged longitudinally around the engine centerline so as to provide an annular gas flow path. The compressor section compresses incoming atmospheric gases that are then mixed with a combustible fuel product and burned in the combustor section to produce a high energy exhaust gas stream. The turbine section extracts power from the exhaust gas stream to drive the compressor section. The exhaust gas stream produces forward thrust as it rearwardly exits the turbine section. Some engines may include a fan section, which is also driven by the turbine section, to produce bypass thrust. Downstream of the turbine section, a military engine may include an augmentor section, or "afterburner", that is operable to selectively increase the thrust. The increase in thrust is produced when fuel is injected into the core exhaust gases downstream of the turbine section and burned with the oxygen contained therein to generate a second combustion.

At start-up, the turbine section of the gas turbine engine has yet to fully provide power. Thus, driving the compressor section may be more challenging than it will generally be at steady state or design conditions. Accordingly, gas turbine engines may include one or more bleed valves to bleed air away from the core flow path during start-up to thereby reduce the load required to drive the compressor.

<CIT>, <CIT> and <CIT> disclose systems of the prior art.

A system for bleeding air from a core flow path of a gas turbine engine according to one aspect of the invention is provided by claim <NUM>.

An optional embodiment includes that the bleed air duct is of a larger diameter than the pressurized air duct.

An optional embodiment includes that the controller is a FADEC.

An optional embodiment includes that the first entrance point is positioned proximate a compressor section of the gas turbine engine.

An optional embodiment includes that the first entrance point is positioned proximate a low pressure compressor section of the gas turbine engine.

An optional embodiment includes that the first entrance point is positioned upstream of a low pressure compressor section of the gas turbine engine.

An optional embodiment includes that the second entrance point is positioned proximate a high pressure compressor section of the gas turbine engine.

An optional embodiment includes that the second entrance point is positioned downstream of a high pressure compressor section of the gas turbine engine.

An optional embodiment includes that the second entrance point is positioned proximate a P3 location within the gas turbine engine.

An optional embodiment includes an anti-ice system in communication with the vortex tube.

A gas turbine engine, according to one aspect of the present invention is provided by claim <NUM>.

A method for starting a gas turbine engine according to one aspect of the present invention is provided by claim <NUM>.

An optional embodiment includes communicating airflow from the vortex tube to an anti-ice system.

It should be appreciated; however, the following description and drawings are intended to be exemplary in nature and non-limiting.

The gas turbine engine <NUM> is disclosed herein as a two-spool turbofan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM>, and a turbine section <NUM>. The fan section <NUM> drives air along a bypass flowpath "B" while the compressor section <NUM> drives air along a core flowpath "C" for compression and communication into the combustor section <NUM>, then expansion through the turbine section <NUM>. Although depicted as a turbofan in the disclosed non-limiting embodiment, it should be appreciated that the concepts described herein may be applied to other engine architectures such as turbojets, turboshafts, and three-spool (plus fan) turbofans.

The engine <NUM> generally includes a low spool <NUM> and a high spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine static structure <NUM> via several bearings <NUM>. The low spool <NUM> generally includes an inner shaft <NUM> that interconnects a fan <NUM>, a low pressure compressor ("LPC") <NUM> and a low pressure turbine ("LPT") <NUM>. The inner shaft <NUM> drives the fan <NUM> directly or through a geared architecture <NUM> that drives the fan <NUM> at a lower speed than the low spool <NUM>. An exemplary reduction transmission is an epicyclic transmission, such as a planetary or star gear system.

The high spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure compressor ("HPC") <NUM> and high pressure turbine ("HPT") <NUM>. A combustor <NUM> is arranged between the high pressure compressor <NUM> and the high pressure turbine <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.

Core airflow is compressed by the LPC <NUM>, then the HPC <NUM>, mixed with the fuel and burned in the combustor <NUM>, then expanded over the HPT <NUM> and the LPT <NUM> which rotationally drive the respective high spool <NUM> and the low spool <NUM> in response to the expansion. The shafts <NUM>, <NUM> are supported at a plurality of points by bearings <NUM> within the static structure <NUM>.

With reference to <FIG>, a system <NUM> for bleeding air from a core flow path C of the gas turbine engine <NUM> includes a bleed valve <NUM> in a bleed duct <NUM>, a pressurized air valve <NUM> in a pressurized air duct <NUM>, and a controller <NUM>. The bleed duct <NUM> receives bleed air from a first entrance point <NUM> to the core flow path C. The pressurized air duct <NUM> receives pressurized air at a pressure greater than that received into the bleed duct <NUM> from a second entrance point <NUM> to the core flow path C downstream of the first entrance point <NUM>.

The core flow path C includes the low pressure compressor ("LPC") <NUM> and the high pressure compressor ("HPC") <NUM>. In various embodiments, the compressor section <NUM> may include one or more compressor stages (each stage including a rotor section and a stator section). The first entrance point <NUM> to the core flow path C is positioned proximate the low pressure compressor ("LPC") <NUM>, and more specifically, proximate a stage 52A of the high pressure compressor ("HPC") <NUM>. In one embodiment, the second entrance point <NUM> is configured to receive the pressurized air proximate a final stage 52N of the high pressure compressor ("HPC") <NUM>. In one embodiment, the HPC <NUM> is a centrifugal compressor and the first entrance point <NUM> is from the shrouded section of the impeller and more specifically aft of the entrance to the impeller inducer, and the second entrance point <NUM> is configured downstream of the impeller and upstream of the centrifugal compressor diffuser section. The centrifugal compressor diffuser section may include swirling vanes or an array of pipes circumferentially surrounding the impeller discharge. In another embodiment, the HPC <NUM> is a multi-stage axial compressor and the first entrance point <NUM> is between the first and second stages, and the second entrance point <NUM> is configured at least one more stage downstream of the second stage and proximate the exit of the one more downstream stage. The locations of the first and second entrance points optimally result in a choked flow condition at the eductor outlet <NUM> within the bleed duct <NUM>. A nearly-choked flow condition means the air pressure at the second entrance point <NUM> is about one and one-half times the air pressure of the first entrance point <NUM>. A choked flow condition means the air pressure at the second entrance point <NUM> is about two times the air pressure of the first entrance point <NUM>. In one embodiment the HPC <NUM> is a multi-stage axial compressor and the first entrance point <NUM> is from between the first and second stages and the second entrance point <NUM> is configured at least two more stages downstream of the second stage and proximate the exit of the fourth stage. In one embodiment the HPC <NUM> is a multi-stage axial compressor and the first entrance point <NUM> is from between the second and third stages and the second entrance point <NUM> is configured at least two more stages downstream of the second stage and proximate the exit of the fifth stage.

The change in the enthalpy of the core flow C imparted by each stage of the multi-stage axial compressor HPC <NUM> is substantially the same. As the temperature of the core flow C increases by passing through each successive stage, the stage pressure ratio of each successively aftward stage of the HPC <NUM> decreases and more stages are needed between the first entrance point <NUM> and the second entrance point <NUM> to result in a choked flow condition at the eductor outlet <NUM> within the bleed duct <NUM>. In another embodiment the HPC <NUM> is a multi-stage axial compressor and the first entrance point <NUM> is from between the second and third stages and the second entrance point <NUM> is configured at least three more stages downstream of the second stage and proximate the exit of the sixth stage. In another embodiment the HPC <NUM> is a multi-stage axial compressor and the first entrance point <NUM> is from between the third and fourth stages and the second entrance point <NUM> is configured at the exit of the last stage 52N of the HPC <NUM>. Locating the first entrance point <NUM> and the second entrance point <NUM> more aftward in the HPC <NUM> increases the density of the air flows in the bleed duct <NUM> and the pressurized air duct <NUM> and reduces the cross-sectional areas of the ducts. Locating the first entrance point <NUM> and the second entrance point <NUM> more aftward in the HPC <NUM> increases the temperature of the air flows in the bleed valve outlet <NUM> and the pressurized valve outlet <NUM>. An increase to the temperature of the air flows reduces the flow Mach number of the air flows and increases heat transfer from the airflows and reduces pressure losses in the bleed duct <NUM> and the pressurized air duct <NUM>. Reducing the flow Mach number of the air flows and increasing heat transfer out from the air flows reduces the frictional and thermodynamic (also known as Rayleigh flow) pressure losses in the bleed duct <NUM> and the pressurized air duct <NUM>.

The second entrance point <NUM> is positioned downstream of the first entrance point <NUM> at a relatively higher pressure location along the core flow path C. In one example, the second entrance point <NUM> may be located proximate an outer annular plenum of the combustor section <NUM> that operates at a pressure of approximately <NUM> psia (<NUM> kPa) and is referred to herein as P3. For example, P1 represents a pressure in front of the fan section <NUM>; P2 represents a pressure at the leading edge of the fan <NUM>; P2. <NUM> represents the pressure aft of the LPC <NUM>; P3 represents the pressure aft of the HPC <NUM>; P4 represents the pressure in the combustion chamber <NUM>; P4. <NUM> represents the pressure between the HPT <NUM> and the LPT <NUM>; and P5 represents the pressure aft of the LPT <NUM>.

The bleed valve <NUM> includes a bleed valve inlet <NUM> and a bleed valve outlet <NUM> in the bleed duct <NUM>. The pressurized air valve <NUM> includes a pressurized inlet <NUM> and a pressurized valve outlet <NUM> in the pressurized air duct <NUM>. The bleed duct <NUM> includes an outlet <NUM> into the bypass flowpath "B" and the pressurized air duct <NUM> includes an eductor outlet <NUM> within the bleed duct <NUM>. One preferred location of bleed valve <NUM> is proximate to the first entrance point <NUM>. Optionally, valve <NUM> in outlet <NUM> can functionally supplement or substitute for bleed valve <NUM>. Together, valve <NUM> is a gate valve with low pressure losses and valve <NUM> is a pressure throttling valve that controls the flow in outlet <NUM>. As shown in <FIG> the pressurized air valve <NUM> is proximate to the second entrance point <NUM>. Optionally, valve <NUM> in outlet <NUM> can be supplementary to bleed valve <NUM>. Together, valve <NUM> is a gate valve with low pressure losses and valve <NUM> is a throttling valve that controls the flow in outlet <NUM>. As shown in <FIG> the pressurized air valve <NUM> can be proximate to the eductor and this placement reduces the length of ducting of the pressurized valve outlet <NUM>. As shown in <FIG>, the pressurized air valve <NUM> must be upstream of the vortex tube <NUM> to regulate the heating and cooling capacity of the vortex tube by changing the pressure at the inlet to the vortex tube. Changing the pressure at the inlet to the vortex tube <NUM> changes the temperature difference between eductor outlet <NUM> and the inlet to the Anti ice unit <NUM>.

The eductor outlet <NUM> is essentially a type of pump which works on the venturi effect to facilitate pumping air through the bleed duct <NUM>. The eductor outlet <NUM> requires only the motive fluid of the higher pressure air from the second entrance point <NUM> for operation. In one example, the bleed duct <NUM> is of a larger diameter (e.g., <NUM>-<NUM> inches or <NUM>-<NUM>) than the diameter the pressurized air duct <NUM> (e.g., <NUM>-<NUM> inches or <NUM>-<NUM>) which facilitates system installation and reduced weight.

The controller <NUM> generally includes a control module <NUM> that executes logic <NUM> (<FIG>) to control operation of the bleed valve <NUM> and the pressurized air valve <NUM> to provide a stable sub-idle initial condition of air flow rate during engine starting. The functions of the logic <NUM> are disclosed in terms of functional block diagrams, and it should be appreciated that these functions may be enacted in either dedicated hardware circuitry or programmed software routines capable of execution in a microprocessor-based electronics control embodiment. In one example, the control module <NUM> may be a portion of a flight control computer, a portion of a Full Authority Digital Engine Control (FADEC), a stand-alone unit, or other system. The control module <NUM> may also communicate with a pressure sensor <NUM> that is used to control opening and closing of the bleed valve <NUM> and the pressurized air valve <NUM>.

With reference to <FIG> and <FIG>, a method <NUM> of starting a gas turbine engine is disclosed. Initially, the method <NUM> includes rotating the core spool (<NUM>), by a starter motor or the like. Next, the control module <NUM> opens the first entrance point to the core flow path C to direct bleed air (<NUM>) from the compressor section to reduce a pressure therein. Next, the control module <NUM> opens the pressurized air valve <NUM> (<NUM>) directing pressurized air from the higher pressure second entrance point <NUM> positioned downstream thereof into the eductor outlet <NUM> to facilitate flow through the bleed duct <NUM> (<NUM>). That is, the system <NUM> for bleeding air induces and entrains a higher rate of bleed flow from the mid-stage of the high pressure compressor and enables a faster start. Finally, the combustor is ignited once the core flow path compressor section and the core flow path turbine section have reached sufficient rotational velocity for start-up.

With reference to <FIG>, the system 100A includes a vortex tube <NUM> in the pressurized air duct <NUM> downstream of the pressurized air valve <NUM>. The vortex tube <NUM> may be a mechanical device that separates the compressed gas into a relatively cold, high density flow <NUM> into the bleed duct <NUM> and a relatively hot, low density flow <NUM> into, for example, an anti-ice system <NUM>.

The system enables a stable sub-idle initial condition of air flow rate during engine starting, especially for starting very high pressure ratio high pressure compressors. The system is also passive and reduces mechanical complexity by avoiding auxiliary rotating turbomachinery and gearing.

Although particular step sequences are shown, described, and claimed, it should be appreciated that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure.

Claim 1:
A system for bleeding air from a core flow path of a gas turbine engine, comprising:
a bleed valve (<NUM>) in a bleed air duct (<NUM>) configured to receive bleed air from a first entrance point (<NUM>) to the core flow path (C) into the bleed air duct (<NUM>);
a pressurized air valve (<NUM>) in a pressurized air duct (<NUM>) configured to receive pressurized air from a second entrance point (<NUM>) to the core flow path (C), the pressurized air at a pressure greater than that received into the first entrance point (<NUM>); and characterised by
a vortex tube (<NUM>) in communication with the pressurized air duct (<NUM>);
an eductor outlet (<NUM>) from the pressurized air duct (<NUM>) located in the bleed air duct (<NUM>); and
a control system (<NUM>) operable to control operation of the bleed valve (<NUM>) and the pressurized air valve (<NUM>).