GB2293415A - High thermal efficiency gas turbine engine - Google Patents

Abstract

The invention may be applied to all classes of gas turbine engine, especially turbo-shaft for electricity generation and automobiles, but also turbo-fan and turbo-jet for aero engines. A very high (86 - 92%) Thermal Efficiency is achieved by manufacturing certain components from Refractory Metals (Cr, Mo, W, melting points 1900 - 3410C) and Carbon-in Carbon, thereby enabling the engine to operate at higher temperatures. In order to prevent oxidation of the refractory components (e.g. ducts 1 and guide vanes 2) means are provided for washing them with fuel to provide a reducing atmosphere. The washing fuel has no available air for combustion until the Second Stage, air-cooled (ODS-Fe-Cr-Al) nozzle guide vane 4, where the excess fuel is burned to provide power to the downstream turbine stages 7, 8. The turbine 7 may be of ODS Ni-Cr. Further details of the materials that may be used and temperatures of operation are given. <IMAGE>

Description

HIGH ThERMAL EFFICIENCY GAS TURBINE ENGINE The present invention relates to gas turbine engines (principally turbo-shaft but also including turbo-fan and turbo-jet aero-engines) to a mode of operating such engines, to components thereof, to alloys suitable for use in such engines and to a method of producing such alloy components.

The proposed gas turbine has a very high Thermal Efficiency (85 - 92%). This efficiency compares with, typically, 37.5 - 45% in current turbines.

It is a fundamental thermodynamic fact that a heat engine becomes more efficient the hotter is its operating temperature (Tmax, the combustor exit temperature). Thus Carnot's Cycle is expressed as follows: Thermal Efficiency = Tmax - Tmin Tmax where T is temperature in degrees K.

Thus, the novel high efficiency is due to the very high Tmax of the heat engine (1600-3200C) but preferably (1800-2177C) The attainment, continuously, of these temperatures using fuel-washed Refractory Metals, in a double high pressure (HP) turbine with complete combustion of the fuel, is the basis of this invention.

The historic Whittle engine had a Thermal Efficiency of only 18% and further progress was dependent on developing high temperature alloys (notably Nickel-Chromium) hardened by precipitates1 solid solution, and more recently oxide-dispersion strengthening (ODS) together with directional grain structure and single crystal components.

Extensive air-cooling (e.g. 35% total air to cool the combustor and 10% for nozzle guide vanes (HGV's) and HP1 rotor blades) has facilitated remarkable increases in gas temperatures. However, the use of so much cooling air is very wasteful since it has to be generated by the compressor.

State-of-the-art oxide dispersion strengthened (ODS) complex Ni-Cr and ODS Fe-Cr-Al have peaked viable metal temperatures at 1 iSOC and 1250-1300C respectively as they are limited by the melting point of the metal matrices.

Refractory Metals have very high melting points, as follows: Element MeltingPoint(C) Density(g/cm ) DBTT(C) * Tungsten 3410 19.3 +100 to +300 Molybdenum 2620 10.2 -100 to +200 Chromium 1900 7.1 +200 to +600 Vanadium 1900 6.2 -3Oto+100 Tantalum 3000 16.6 -200 Niobium 2470 8.6 -125 ' Ductile / Brittle Transition Temperature ODS-Tungsten - La2 03 and TZM - Molybdenum (hardened with carbides of titanium and zirconium) are commercially available and applications include: electric light filaments, isothermal forging dies, high temperature furnace hardware, diecasting tools, radiation shields, heat insulation, heating elements, etc.

However, service applications have to be under vacuum (eg ODS-Tungsten in the domestic light-bulb at 2700C) or under controlled furnace atmospheres (eg H2, N2, Ar). This is because the Refractory Metals are very susceptible to oxidation at comparatively low temperatures (erg296 - 1000C).

This invention maintains a fuel-rich reducing atmosphere in the gas turbine at Tmax temperatures up to 3200C but preferably 2127C (stoichliometric for Kerosenc) with ODS Tungsten or 1800-1 900C with TZM-Molybdenum and the very poor oxidation resistance of the Refractory Metals is overcome allowing the very high temperature potential to be realised.

The reduced Tmax, necessary for TZM-Molybdenum imposes less thermal stress on ancillary equipment of the engine while maintaining a Thermal Efficiency of 86% with a much lower density (10.2 gXcm3) than ODS-Tungsten (19.3 g/cm3). This opens the way to aero-engine and automotive applications. The lower heat content enables a 2-shaft design as the lower heat energy can be absorbed by the twin HP turbines. This simplifies the engineering of the turbine. In a turbo-shaft gas turbine it is necessary for efficiency to completely convert the heat engine fuel to power for the main shaft and the compressor, whereas for aero-engines any excess heat in the jet-stream is utilised as thrust.

Thus one type of embodiment of the invention provides an extremely high efficiency gas turbine engine (typically 86 - 92% thermal efficient) due to the use of fuel washing of Refractory Metals and Carbon-in-Carbon.

Fuel-washing provides a controlled, reducing atmosphere from Turbine Entry Ducts in the combustor (Tmax, (1) on the figure) through the First Stage nozzle guide vanes (2) and the First Stage high pressure (HP1) rotor blades (3). The fuel for washing is fed through the skins of the doubleskinned combustor (5 - 6) thereby cooling the ODS-Fe-Cr-AI inner skin to the surface of the TED's via small exit holes, thereby cooling T4 to 1800-1900C to accommodate TZM-Mo or ODS-W. The bulk of the fuel is burned stoichiometrically (2127C) for Kerosene, 2177C for hydrogen, so that the washing-fuel (eg 5-15% of total fuel) is in excess of stoichiometric and cannot burn since there is no available oxygen.It therefore provides the necessary controlled reducing atmosphere, within the core of the engine, for Refractory Metal temperatures of (1600-3200C) with corresponding Thermal Efficiency. Of course, a more moderate Tmax of (1800-2177C) is preferred, with only a 4% Efficiency penalty (eg 86%).

A second type of embodiment of the invention provides a novel technique to utilise the excess washing and cooling fuel (which, if wasted would negate the improved Thermal Efficiency from the high Tmax). The second stage NGV (4 on the figiure) preferably in ODS-Fe-Cr-AI is air-cooled, since oxidation resistance to 125011300C of this alloy is outstanding. The fuelrich reducing atmosphere ignites (using a fuel-burning pilot light for safety) with this excess to stoichiometric NGV-exit air, thereby (as a twin HP turbine) generating heat and power to the Second Stage rotor blade HP2 (7) on the figure. (7) is in ODS-complex Ni-Cr directionally solidified or single crystal complex Ni-Cr.

Preferably the Second Stage HP2 turbine drives the compressor while, unusually, the High Tech (1600-1800C) First Stage HP1 turbine drives the main shaft, rather than being wasted by driving the compressor.

A third type of embodiment of the invention is a flexibility of the Efficient Gas Turbine between electricity generation and heat from the exhaust.

A two shaft design is preferable to three shaft because engineering is easier. However, it is not always possible to absorb all of the heat energy with only two turbines in the turbo-shaft engine. Thus it is best to use the semi-domestic 100-2000KW engine in a total electrical power (preferably hydrogen) environment eg on an isolated fish farm with no access to a grid system, so that any excess heat at the tail-end of the jet-pipe can be fed into a water (preferably waste water) heating system consisting of Cupro-Nickel tubes to heat the fishbreeding tanks from eg 6C to 10C. This warmth stimulates the breeding process. Buildings on the fish farm can also be heated from this source. Alternatively the excess heat can be used to drive a steam turbine.

A fourth type of embodiment of the invention provides a Twin HP Gas Turbine Engine with little or no air-cooling of the combustor barrel, turbine entry ducts, First Stage NGV's or First Stage rotor blades. The savings in air are, typically, up to 35% from the combustor and up to 10% from NGV's and rotor blades.

35% combustor peripheral cooling air results in a profile in the HP1 rotor blade with a hot spot at the centre of the blade, which limits the temperature ability of the blade. Elimination, or substantial reduction, of this peripheral cooling air (eg by the use of ODS-Fe-Cr-Al combustor liners) has been shown to achieve 1350C with a very flat temperature profile.

ODS-Tungsten or TZM-Molybdenum increase this temperature capability in the TED's to (eg) 1800-2177C.

These substantial and unprecedented savings in cooling air provide efficiency in addition to the fundamental 85-92% thermodynamic efficiency. This total efficiency is cashed as a quantum leap forward in low fuel consumption with cut-back in C02 emissions and global warming.

Kerosene is the preferred fuel because it is well-established but hydrogen may be used where hydrocarbons are not permissible eg because of carburisation above 900-1100C of Refractory Metals Mo and W. If hydrogen is used as fuel (eg in a small turbo-shaft engine in power generation and automobiles) C02 emissions and global warming are zero.

The low cooling air requirement puts a very light load on the compressor and a low pressure ratio (PR) of around 15% is viable and air-heating is minimal with beneficial spin-off for efficiency.

At 15% PR, compressor stall is not a problem and the usual expensive precautions of three shaft design to allow two shifts for a split compressor or variable angle compressor guide vanes (two shaft design) are unnecessary.

A fifth type of embodiment of the invention provides (in an optional three shaft design) a low pressure (LP) turbine stage (8 on the figure) to extract the last of the heat energy for turboshaft applications. Ideally all of the power is absorbed by the Twin HP1 and HP2 Turbines at a Tmax (TZM-Mo) of 1800-1900C. For aero-engine applications, where excess heat is used as thrust a two shaft design is always used.

A sixth type of embodiment of the invention provides for a fuel-washed Twin HP Gas Turbine Engine, utilising Refractory Metal Molybdenum alloys in combustor TED's and First Stage NGV's to give a Tmax of 1800-1900C and a Thermal Efficiency of 86% with preferably a two shaft design, although three shaft may be necessary. The preferred Molybdenum alloy is TZM (0.5 Ti - 0.O8Zr) but Mo-0.5Ti, Mo-30W, Mo-OSZr, Mo-0.05Zr, Mo-1.3STi - 0.15Zr - 0.15C (TZC) or potassium silicate-doped Molybdenum may be used.

The relatively low-density (10.2 g/cm3) in these static components, makes aero-engine applications (turbo-fan or turbo-jet) viable. Like many body-centred cubic Refractory Metals, especially tungsten and chromium, TZM Mo has a higher than ambient Ductile / Brittle Transition Temperature (DBItI) of -100 to +200C. TZM-Mo is perfectly ductile from 2002620C.

A seventh type of embodiment provides for a fuel-washed Twin HP Gas Turbine Engine utilising ODS- Refractory Metal Tungsten, as an alternative to TZM- Molybdenum in combustor TED's and First Stage NGV's to give a significantly higher Tmax of 1600-3200C but preferably 1800-2177C and a Thermal Efficiency of up to 92% with two or three shaft design. ODS-Tungsten DBTT of 1 00-300C has been perfectly adequate in domestic lightbulb filaments since it was patented in 1908.

The oxide used in ODS-W is preferably La2 03 or any rare earth oxide, MgO, A12 03 or Th02.

ODS-W offers improved oxidation resistance over Molybdenum (TZM or otherwise), higher strenth at (eg) 1800C than Mo, and higher temperature capability. Its very high density (19.3 glum3) is perfectly satisfactory in the non-rotating components, combustor TED's and First Stage NGV's in static turbo-shaft Gas Turbines, such as are used in electricity generation.

However the dead-weight is too high for aero applications. ODS-Tungsten is ductile from 300-3410C.

Continuous solid solubility of Vanadium and Chromium in Tungsten enables alloying (eg 1530%) to substantially reduce the density of a Tungsten component.

An eighth type of embodiment of the invention comprises a First Stage HP rotor blade (preferably in ODS-Chromium) which, with no cooling, will withstand Tmax gas of 1800C (TZM-Mo for combustor TED's and First Stage NGV;s) with a blade metal temperature of 1650C. Tmax of 1900C (upper limit of TZM-Mo) gives a blade metal temperature of 1750C.

Any cooling of the HP1 blade must be by fuel, so as not to upset the fuel-rich atmosphere. if blade fuel-cooling of 230C is used, 2127C (stoichiometric temperature for Kerosene) utilising ODS-Tungsten, a blade metal temperature of 1750C is attained (these figures assume a modest temperature drop of 150C between combustor and HP1 Turbine blade).

Any cooling fuel is burned off at the second stage NGV (air-cooled to end the reducing atmosphere) together with the washing fuel. ODS-Chromium has a suitably low density (7.1 glum3) for rotating components (lower density than a Nimonic alloy).

A rare-earth oxide such as Y2 03 imparts outstanding nitridation resistance to a Chromium base alloy. Early Australian (Alloy H) or USA alloys would be totally penetrated by brittle nitride (eg 15cm dia bar) after exposure to a moving air-stream at 1100-1400C for eg, 24 - 48 hours.

A usable nitridation resistant Chromium alloy for Gas Turbines has long been sought and nitridation resistance by virtue of a very fine dispersion of rare-earth oxide is now a reality.

Modest high temperature strength is imparted by oxide and carbide dispersions, which also punch out dislocations to ductilise the matrix, giving a DBTI of 200C in slow bend and 600C in impact. Thus the DBTT of body-centred-cubic ODS-Chromium is of the same order as ODS-Tungsten and TZM-Molybdenum. Moreover, at 600C, the Cr-Y2 03 bar stopped the hammer of the Izod impact machine i.e. was perfectly ductile from 600-1800C.

Low temperature machining of Cr-Y2 03 is possible in compression (eg centreless grinding) or by comparatively new techniques such as electro-chemical or laser machining, but not electro-discharge (EDM).

The way forward to high temperature operation of Refractory Metals, such as Chromium, Molybdenum or Tungsten is to heat the gas turbine gently to above the DB1T (eg 500C) and then to utilise the properties at temperatures many hundreds of degrees above DBTT, where the ductility is perfectly sound.

A ninth type of embodiment provides a production route for high purity (with respect to residual 02 and N2 levels) ODS-Chromium, using high purity hydrogen to reduce / sinter a forging blank and finally to press-forge the blank to near net rotor blade shape.

A tenth type of embodiment provides for fuel-washed Refractory Metal ODS-Chromium in First Stage (HP) rotor blades: (i) Alloyed with oxide only Cr- (0.5-2.5)% Y2 03 (or other rare-earth oxide) MgO A12 03 or Th02 Preferred low-alloy composition: Cr- 1.36% rare-earth oxide. (Y2 03 or La2 03) (ii) Alloyed as follows: As (i) but in addition: (2-7)% Al for improved oxidation resistance (0.5-2)% V to enhance nitridation resistance (0.5-2)% (Ta, Ti, Nb, Zr) carbides for ductility (0.5-10)% W for solid solution hardening Preferred highly-alloyed composition: Cr- 1.36Y2 03 - 4.5 Al - 4.5W- 1 V - 1.5 TaC KEY TO ENGINE DESCRIPTION 1. ODS-Tungsten, TZM-Mo or C-in-C Turbine Entry Ducts.

2. First Stage NGV in ODS-Tungsten, TZM-Mo, or C-in-C.

3. First Stage Rotor Blade (eg 230C fuel cooled) in ODS-Chromium, C-in-C or other suitable blade material.

4. Second Stage air-cooled NGV in ODS-FcCr-A1 featuring burning of excess fuel from fuel-washing and fuel cooling.

5. ODS-Fe-Cr-A1 or C-in-C sheet Combustor Liners.

6. Outer conventional weldable sheet eg C263, PK33, HS188, Nimonic 86.

7. Second Stage Rotor Blade in eg ODS-Ni-Cr (MA6000), directionally solidified Ni-Cr or single crystal Ni-Cr.

8. One or two rows of conventional Nimonic rotor blades (eg Nimonic 115) in three shaft design. Deleted in preferred two shaft engine.

Claims (8)

1. A gas turbine engine having components which comprise refractory materials1 the engine being adapted to operate so that at least parts of the components are raised to temperatures at which they are prone to oxidation damage; and wherein there are means for effecting fuel-washing of at least the damageprone parts so that they are protected by a reducing atmosphere.

2. An engine according to Claim 1, wherein the Refractory materials are selected from materials based on Chromium, Molybdenum, Tungsten and Carbon-in-Carbon.

3. An engine according to Claims 1 and 2 adapted to operate with combustor exit temperatures in the range 1600-3200C.

4. An engine according to any preceding Claim wherein the fuel-washing is applied to achieve a reducing atmosphere from combustor turbine entry ducts (TED's) through First Stage nozzle guide vanes (NGV's) and High Pressure (HP1) rotor blades.

5. An engine, as claimed in any preceding Claim where fuel used for washing and cooling is burned efficiently at the Second Stage excess air-cooled NGV's.

6. An engine, as claimed in any preceding Claim and having an ODS-FeCr-Al skin and surface of the ODS-W, TZM-Mo or Carbon-in-Carbon components which surface is fuel-cooled to reduce TED temperature to 1800-1900C.

7. An engine, as claimed in any preceding Claim with a nitridation-resistant uncooled or fuel-cooled HPl Twin Turbine rotor blade in ODS-Cr (preferably Cr-1.36 Y2 03 4.5Al - 4.5W - IV - .5TaC) at 1600-1800C metal temperature.

8. A process of combining necessary high purity hydrogen reduction (of N2 and 02) of ODS-Cr of the composition specified in Claim 7, with forging preform sintering to forge to high density near-net rotor blade (HP1) shape.