GB2307008A - Gas turbine engine with two stage combustion - Google Patents

Abstract

In a gas turbine engine partial combustion of a fuel at E, at a high temperature but with a deficiency of air, yields a gas rich in carbon monoxide and hydrogen. This allows the use of turbine blade and casing materials of metals and their alloys with high melting points and strengths at high temperatures, which would corrode in the oxidative conditions encountered in normal gas turbine combusters. The rich gas after expansion to an intermediate pressure in a turbine F to provide work, is burnt to completion at G with further excess air and expanded again in a turbine H, providing further work. The invention may be used in either turbo-jet engines or non-regenerative gas turbines. In both cases substantial gains in thermal efficiency may be realised compared with conventional engines.

Description

SPECIFICATION: TURBINES.

This invention relates to improvements in turbo-jet and non-regenerative gas turbine engines which lead to greater efficiency.

It is well known that fuel efficiency can be increased in engines of these types by raising the combustion temperature of the fuel/air mixture supplied with the aid of rises in operating pressure. This is based on well known thermodynamic principles. A limit is set to efficiency gained by the ability of the materials available for the construction of the expansion turbine blades to withstand this temperature, both with respect to strength and corrosion and oxidation in the hot gases. Much research work has been conducted over past years on various metal alloy combinations and ceramics to improve high temperature strength and comosionloxidation resistance, which has also included work on non-corrosion resistant strong alloys covered with corrosion resistant coatings.Latterly efficiency gains derived in this way have only been small, as the materials which have the necessary properties operate at temperatures closer to their melting points, and thus lose strength rapidly. The invention described in this patent provides a method of surmounting these limitations and achieving significant efficiency gains.

The present invention in its broadest embodiment operates the engine with two combusion chambers, the first at higher pressure than the second. In the first chamber, air at high pressure is combusted at a specially designed bumer with an excess of fuel over that necessary for complete combustion to carbon dioxide and water. The manner of operation is the same as in a synthesis gas producer. The hot gas thus produced is rich in carbon monoxide and hydrogen, and contains only small amounts of carbon dioxide and steam. In the second combustion chamber, at lower pressure, after expansion of the partially combusted gas produced in the first combustion through a turbine, further air at the lower pressure is injected and the partially combusted gases from the turbine are bumt to completion.This completely burned mixture is then expanded through a second, lower pressure, turbine. Work provided by each expansion turbine is used to drive the necessary air compressors, and surplus energy, as in conventional practice, is used either as hot gas discharged via a nozzle in turbojets or as surplus shaft horsepower to drive machinery in a gas turbine. If desired all compressors and turbines can be on one shaft, air at lower pressure being taken off at an intermediate stage of the air compressor, bypassing the first combustion and expansion turbine,and being introduced to the second combustion chamber.

The key to the invention lies in the operation of the first combuster. As has been noted, by restricting air this is operated as a high pressure gasifier. Because the gas produced is rich in carbon monoxide and hydrogen, the first expansion turbine blades can be made of metals and their alloys whose use is normally precluded by their susceptibility to oxidative corrosion.

Such metals as molybdenum, tungsten and rhenium and their alloys with each other and other metals, all of which have high melting points and high strength at high temperatures, may now be used. This in tum permits very substantial temperature rises in this turbine.

Thus for example using thermodynamic data for these metals and their oxides taken from Kubaschewski and Alcock, Intemational Series on Materials Science and Technology; Volume 24; Metallurgical Thermochemistry; Fifth Edition published by Pergammon Press,it can be shown that molybdenum cannot form an oxide at any temperature between 500 degrees K and 1850 degrees K, provided the carbon dioxide carbon monoxide ratio is less than 0.25. In the case of rhenium the ratio is higher because of the lower stability of its oxides. For tungsten the limits are slightly tighter: no oxide will form between 500 degrees K and 1850 degrees K. provided the ratio does not exceed 0.055 (the ratio rises with temperature and is 0.072 at 600 degrees K and 0.184 at 1850 degrees K .).The composition of the synthesis gas produced in the first combustion must of course meet these ratio criteria to prevent oxide formation.

There are of course difficulties in the use of these materials of construction. Tungsten for example passes through a brittle transition at 473 degrees K, but this can be eliminated by alloying with some molybdenum, rhenium or both. Working of these metals also requires special techniques by virtue of their high melting points, particularly unalloyed tungsten.

Rhenium, perhaps the best choice, is expensive. Rhenium and tungsten are very dense, giving heavier engines and bigger blade stresses at a given turbine speed. These problems are to some extent offset by the smaller diameters of combusters and turbines operating at higher pressures.

Molybdenum and tungsten form carbides by "gettering" carbon from carbon monoxide unless sufficient carbon dioxide is present. This is, of course, the reverse of the ratio of these gases required to prevent oxidation, but there are ratios at which nether will occur.

Thermodynamic calculations based largely on data from Kubaschewski but also from E.K.

Storms, Refractory Carbides, Academic Press 1967 and Molybdenum-Physico-Chemical Properties of its Alloys, Atomic Energy Review, Special Issue, No.7, 1980, show that the carbide of molybdenum could potentially form at all desirable operating temperatures but the highest (the carbide is less stable in carbon monoxide at higher temperatures and, besides, more carbon dioxide is present in the first stage combuster when a higher temperature is required at the same operating pressure). This is not so with tungsten, although certain operating conditions have to be avoided. The problem arises after the first expansion, even though the pressure is lower which hinders carbide formation. The gas cools, carbide becomes more stable at lower temperatures, and this effect is greater than that due to pressure fall.Some of the lower temperature, higher pressure cases listed subsequently in Table 1 are on the borderline of carbide formation. Here, if desired, the temperatures are low enough to use nickel alloys instead. Carbide could of course form as the engine warms up after starting. Any formed (and this may be limited by slow reaction speeds at lower temperatures) will revert to metal at operating temperature. There is also scope for warming the engines using slightly higher air fuel ratios in the first stage combuster, thus depressing the carbon monoxide carbon dioxide ratio and, incidentally, helping to prevent the formation of carbon itself.

The use of molybdenum cannot be ruled out because it tends to form a carbide under likely operating conditions. The carbide is not particularly stable and judicious alloying may prevent it (Also alloys generally have a negative free energy of formation; that is they are more stable than the parent elements. Even if the sole effect is a reduction in the chemical activity according to Raouits Law, that may suffice. This is particularly the case with molybdenum because there are two atoms present in its carbide; thus a molybdenum alloy mole fraction of say 0.6 would reduce the activity of the molybdenum in it to around 0.36).

Molybdenum can also form a nitride, just stable below about 1300 degrees K in the combustion synthesis gases at higher total pressures. As will be seen, operating temperatures will usually be above this in the first stage expansion at all pressures and again suitable alloying could inhibit its formation. Generally also the rate of formation of the nitride is said to be very slow below 1400 degrees K.

When the invention is to be used in aircraft engines, the lower pressures at operating altitudes (1820% of those at ground level) will raise the ratio of monoxide to dioxide at which carbide formation can occur with these metals and their alloys because these reaction equilibria are pressure dependent. Pure molybdenum can now be used at first turbine inlet temperatures, but carbide would form at the turbine exit temperatures. Tungsten is acceptable for use under all conditions. Similarly molybdenum nitride will not form at all operating conditions at altitude.

Finally it is worth noting that silicon carbide will not oxidise to silica in the first stage combuster gases at any temperature above 1000 degrees K. This material has been used in carbide type material blade tests in normal oxidizing conditions, even though it is too brittle.

Silica forms and it has generally been hoped that an impervious layer would control sub surface silicon carbide oxidation. The use of this invention would prevent silica formation altogether.

The nature of the invention will be more readily understood by reference to Figure 1, which relates to its use in a turbo jet engine in an aeroplane in which one shaft is used for simplicity. Different arrangements are possible elsewhere, ranging from the use of overlapping anular shafts to permit differing compressor and turbine speeds, to complete separation of all stages in land based applications. Modifications to the drawing are required, such as the omission of a nozzle, if the invention is to be used as a non regenerative gas turbine, but these will be obvious to those skilled in the art. Refening to the Figure, air is injected via a conventional infuser A and compressed by compressor blades B to the pressure of the second stage combuster G.A portion of this air is taken into a further stage of compression by compressor C , the volume being controlled by the diameter of the casing D and first staging of C and by louvres or valves as necessary. After compression, the air partially bums fuel injected as at K into combustion chamber E to provide the gas mixture rich in carbon monoxide and hydrogen. Careful attention will be necessary to the design of this bumer to eliminate or minimise fine carbon particle formation. The design should draw on techniques used in high pressure oil gasification, such as the so called toroidal bumer used in the Shell gasifier, or a modification used in the Texaco oil gasifier.The fuel, which may be gaseous, liquid or solid, may be preheated, say by indirect heat exchange with the hot compressed air or another heat source (exhaust gases for example in the case of a gas turbine applications) and, if liquid, may be vaporised by the same means before injection.

Liquid fuels such as paraffin will be used in air transport applications. Careful control of the air flow and fuel flow from the fuel pumps is required to achieve the correct gasification ratio.

In land applications, water or steam can be added to aid control of carbon formation.

After partial combustion, expansion takes place through turbine F, which provides work to drive compressor C and may leave a surplus of energy to drive compressor B. After expansion of partial combustion gases, the bypass air contained in casing L is supplied to combust the rich gas completely in bumer G, which more closely represents bumer designs in conventional engines, except that flame stability may be enhanced because the fuel (products of the earlier partial combustion) is both gaseous and hot. All the products of combustion are then expanded through turbine H, which provides the balance of energy needed to drive compressor B. The hot tail gases from H are passed via a nozzle I to atmosphere to propel the aeroplane. The dotted shaft J would be present to take off power, and the nozzle would be absent in any gas turbine operation.

Casing D must be of sufficient strength at its narrowest diameter to contain the maximum pressure difference been E and G, but usually the diameter at E is quite small. Similarly L, of much larger diameter, must withstand the pressure at G. The design of casing D requires careful attention. A composite construction will best serve. In this a thin nor-load bearing inner shell is made up of high temperature metal or alloy capable of withstanding corrosion and the high temperature gases in E, a material similar to but not necessarily the same as is used in the hottest blades in turbine F. A gas tight outer shell of larger diameter thicker material such as nickel alloy, surrounds the casing. The space in the anulus deliberately left between the casings, is carefully packed with high temperature insulating material.One such material might be tightly packed magnesium oxide powder. Partial combustion products are allowed to leak at points through the inner shell to fill the spaces between the magnesia powder thus preventing oxidation of the inner shell and the inner side of the outer shell.. The outer nickel alloy shell is kept at a sufficiently low temperture by an adequate thickness of insulating material so as to prevent oxidation by air on its outer side, and give it adequate tensile properties to minimise the metal thickness required to contain the differential pressure between E and G. The passage of relatively cool air around D, the outside of the casing, allows heat transfer outwards to it, which aids the establishment of the correct temperature profile through the multilayered wall. Fins can be fitted to the outer shell if necessary to aid this process.The wall thickness of casing D can, if desired, vary along its length and be thinner where temperatures and pressure differentials are lower, subject to the constraints imposed by increasing diameter.

For a more quantitative understanding of the benefits to be derived from the present invention with respect to fuel efficiency, a series of calculations have been made, whose results are illustrated in Tables 1 and 2. Table 1 gives data for use of the invention in turbo jet applications and Table 2 for its use in non regenerative gas turbine applications.

Refening to Table 1, a comparison is first m. between applications of the invention over a range of temperatures and pressures and da or a conventional turbo-jet engine illustrated on page 787 et seq., of Chemical Process P oles; Part 2; Thermodynamics; by O.A.

Hougen, K.M. Watson and R.A. Ragatz; Se. rdHion, publishers J. Wiley and Sons Inc.

The operating conditions of the conventiona ample given in Chemical Process Principles are not by modem standards particularly arr , the turbine inlet temperature and pressure being only 1144 degrees K (1600 degrees F 8.68 Atm. Nevertheless the comparison is useful because the material in Chemical Pn ) Princioles shows the method of calculation adopted. This result is shown in line no. 1 s Table. Column 8 nl, is the thermal efficiency, column 9, nP, is the propulsion ency and column 10 , the overall efficiency, is the product of these two.Elsewhere in the te, at Line 6, the Chemical Process Principles calculations are uprated to inlet C lanes of 1311 degrees K (1900 degrees F) and 21.82 Atm., which are very arduous ino ven by present day standards.

The fuel used in the calculations which are ared with those in Chemical Process Principles to illustrate the invention, simulat plcal paraffin. It has the formula C 7.2 H13 6 (empirical formula CH 1.89) which ig: trace elements. Its net heat of combustion is 18,572 BTU/Ib.(43.1 KJoules/grm.), its he; formation from its elements 710 BTUllb.(1.647KJoules/grm.), and its heat of isation 131 BTUAb.(0.3039KJoules/grm.).

Dry air is assumed, containing 20.9% oxyge- 79.1% inerts (nitrogen). In calculating the data, use has been made of air tables and t zies at pages 669 to 671 of the above book.

Other data on heats of combustion, specific i, etc., and also enthalpies when outside the range of pages 669 to 671, have been take Technical Data on Fuel; Editor H.M.

Spiers, Fifth Edition; Publisher: The British onal Committee, World Power Conference.

To simplify the calculations at very little los. accuracy, fuel rich first stage combustions at pre-chosen temperatures of 1750 degrees F 1850 degrees K are assumed to produce nether methane nor carbon (this is justifiec e very high temperatures in the first combustion).Dissociation of water, hydmt: carbon monoxide and carbon dioxide are assumed negligible. Also the water gas sh uilibrium in the product gases is assumed to hold in the combustion, but the equilibrium ant used is some 50% higher at each temperature than quoted by Spiers, becaus ification is known to tend to overproduce carbon monoxide and water at the expense carbon dioxide and hydrogen (thus at 1850 degrees K, a value of 6.7 is assumed for tz instant rather than 4.3).The shift is assumed to be frozen immediately after cow stion and is not readjusted when cooling takes place through expansion in the turbine F e gure 1, although some small adjustment is likely in practice (failure to make this adjust tt Is in fact slightly prejudicial to the efficiency of operation of the invention. If the shift dl "just, first stage expander exit temperatures would be some10 degrees K higher).

All examples in both Tables use the same hinery efficiencies as those in Chemical Process Principles. These are 0,88, 0.83 3, 0.86, and 0.85 for the diffuser ram, compressor, combustion, turbine and nozz. fficiencies respectively. As in the book, fuel is injected to the bumer without preheat or vd isation. (Complete fuel vaporisation, when this is possible by heat exchange with, say, the : compressed air supplied, is an altemative not considered. In practice such techniques co- discourage any tendency to form carbon particles at the high pressure bumer.) In Table 1, column 3 is the combustion terT Erature of the first stage.A desired temperature is assumed, either 1750 degre s K or 1850 degrees K, and the fuel air ratio, column 2, to give a heat balance at this temperature is calculated for an assumed operating pressure in Atmospheres shown in column 4. Column 5 shows the temperature in degrees Kelvin of the rich gas after expansion to the pressure shown in Atmospheres in column 6.

Column 7 shows the temperature achieved after combustion of the rich gas with further air at the pressure shown in column 6. When only conventional one stage combustion is used, column 3 merely shows the temperature achieved on combustion at the pressure shown in column 4. Thus Line 1, the conventional case from Chemical Process Principles, has 1144 degrees K in column 3 and 8.68 atmospheres in column 4. In all cases, in all examples, the overall air to fuel ratio supplied is 33.0 Ib.(grm.) moles of air per Ib.(gnm.) mole of fuel, the same ratio as the ratio used in the Chemical Process Principles examples. The invention is not of course limited to this ratio alone, but choice of this fixed ratio allows efficiency comparisons under similar conditions.

By way of illustrating the overall reaction taking place at each stage of the combustion process, the conditions chosen for Line 2 of the Table are used. Here, fuel is bumed in the first stage at an air fuel molar ratio of 3.13 with air which has been compressed to 34.72 Atm.

Taking into account the heat content of this air and assuming no fuel preheat, the reaction which represents this is: 0.791 N2 + 0.209 02 + 0.32 Cm1 .89 = 0.30 CO + 0.078 H20 + 0.224 H2 + 0.02 CO2+ 0.791 N2 (Water gas shift at 1750 degrees K assumed to be 5.2) which yields a heat balance when the temperature is 1750 degrees K. In the second stage, further compressed air at 6.8 Atm. is bumed with the rich gas from stage 1 as shown by the following reaction: 0.30 CO + 0.078 H20 + 0.224 H2 + 0.02 CO2 + 0.791 N2 +7.562 N2+ 1.998 02 = 0.32 CO2 + 0.302 H2O + 8.353 N2 +1.73602.

Here the quantity of secondary air is chosen to bring the overall air fuel ratio up to 33 molar.

This sets the temperature given by the heat balance in this combuster at 1057 degrees K.

With a 18,572 BTUflb.(43.1 KJoules/grm.) net heat of combustion, this quantity of fuel gives the same heat supply of 270.2 BTUflb.(0.627KJoules'grm.) of air supplied as in the example in Chemical Process Principles.

Inspection of Table 1 shows that the overall efficiency of the Chemical Process Principles example shown in Line 1, column 10, is 0.163. Compared with this it can be seen that the overall efficiencies for the two stage expanders with first stage combustion to rich gas, column 10, Lines 2, 3, 4, 5, 7 and 8 inclusive, are 0.169, 0.177, 0.167, 0.195, 0.171 and 0.198 respectively. Compared with the Chemical Process Principles example, these represent efficiency gains of 3.7%, 8.6%, 2.5%, 19.6%, 4.9% and 21.5% respectively.

The highly uprated single stage engine with a blade temperature of 1311 degrees K and inlet pressure of 21.82 Atm., Line 6, has an overall efficiency of 0.186. This is probably at the very limit of temperature for conventional operations. Only the more arduous conditions of two stage operation can improve significantly on this efficiency. Thus the two examples with first stage inlet pressures of 77.87 Atm, Lines 5 and 8, with efficiencies of 0.195 and 0.198, are the only two to give a better performance. Efficiency gains are 4.8% and 6.5% respectively.

Generally examination of Table 1 shows the outcomes which would be expected thermodynamically. Efficiency is little affected by higher first stage temperature unless the temperature rise is achieved by introducing higher enthalpy first stage air by higher compression (compare Lines 2 and 7 and lines 5 and 8 for temperature effect; Lines 4 and 5 for pressure effect). Decreases in air fuel ratio and increases in pressure both increase the potential for carbide formation; a reaction which, as noted earlier, is pressure dependent.

For example for the reaction: W+2CO=WC+ CO2 P(CO)2 : P(CO2) actual equals 500 Atm. for the example in Line 5 and only 223 Atm. for the example in Line 4. On the other hand, comparing the examples of Line 5 at 1750 degrees K and Line 8 at 1850 degrees K., the rise in temperature raises the equilibrium constant for tungsten carbide formation from 1000 Atm. to 2100 Atm. (The air fuel ratio also has to increase between the examples in Lines 5 and 8 in order to raise the temperature. The effect on the actual carbon monoxide carbon dioxide ratio is quite small however, being 500 Atm. at 1750 degrees K and 458 at 1850 degrees K. ). In all cases, as remarked earlier, the monoxide dioxide ratio is not high enough to permit carbide formation.

Chemical Process Principles illustrates the well known improvement in efficiency from operating a one stage engine at 12,195 metres altitude. The data from the book are reproduced in Line 9 of Table 1. The overall efficiency is 0.171. This is an improvement of 4.9% compared to sea level. Line 10 shows operation of the two stage engine with first stage combustion set at 1750 degrees K. Efficiency is 0.204 which is 19.3% better than the single stage engine at altitude.

Table 2 shows the efficiency gains which may be realised by using two stage combustion in non- regenerative gas turbines. The gas turbines are assumed to be fitted to propeller driven aeroplanes operating at sea level and to have the same air inlet conditions as the turbo jet engines of Table 1 taken from Chemical Process Principles (15 degrees C, 1 Atm.

before diffuser: 51 degrees C, 1.446 Atm. after diffuser). Line 1 shows the conventional one stage engine. The first stage pressure is chosen at 6.8 Atm. which, at the set fuel ratio used throughout Tables 1 and 2, gives a combustion temperature of 1106 degrees K and a thermal efficiency of 31.48%. Two two stage examples are given. Line 2, has a first stage combustion temperature of 1750 degrees K, which requires a first stage fuel ratio of 3.13 moles per mole at an operating pressure of 34.72 Atm. The second stage operates at the same 6.8 Atm. pressure as the single stage example. Efficiency is 34.75%: 10.4% better than the single stage machine. Line 3 gives similar data at the same first stage air fuel ratio of 3.13, but this time with a first stage combustion temperature of 1850 degrees K.This higher temperature at the same fuel air ratio requires the inlet pressure to rise to 56.87 Atm.

Efficiency is 36.16%; 14.9% better than the single stage operation at the same overall air fuel ratio.

TABLE 1 Efficiency Improvements from Two Stage Combustion in Turbo Jet Engines. Overall Air/Fuel Ratio 33.0 Molar throughout ( Same as 270 btu. per Ib. Air in Hougen, Watson and Ragatz Examples.) nl Intemal Efficiency, nP Propulsion Efficiency, nl x nP Overall Efficiency.

1 2 3 4 5 6 7 8 9 10 Line 1st. Stage 1st. Stage 1st. Stage Rich Gas Second Temp.after nl nP nl x nP No. Air/Fuel Combust Pressure Temp after Stage buming Ratio.Mole Temp. Atmos. Expansion Pressure Rich Gas.

per Mole Deg. K Deg. K Atmos. Deg. K 1 33 1144 8.68 NA NA NA 0.281 0.579 0.163 2 3.13 1750 34.72 1272 6.8 1057 0.296 0.57 0.169 3 3.03 1750 52.08 1239 8.74 1094 0.316 0.56 0.177 4 3.03 1750 52.08 1475 21.82 1274 0.294 0.568 0.167 5 2.86 1750 77.87 1360 21.82 1274 0.365 0.535 0.195 6 33 1311 21.82 NA NA NA 0.34 0.547 0.186 7 3.33 1850 34.72 1355 6.8 1053 0.301 0.568 0.171 8 294 1850 77.87 1448 2w2 1275 0.373 0.531 0.198 9 33 NA NA NA NA NA 0.329 0.522 0.171 10 3.23 1750 10.79 1239 1.79 1026 0.381 0.534 0.204 All examples at sea level, 15deg., C., and 1 Atmos., except lines 9 and 10 which are at 12195 m., -56deg., C., and 0.185 Atmos. external pressure. These two lines are shown in bold type.

TABLE 2 Efficiency Improvements from Two Stage Combustion in Non-regenerative Gas Turbines. Overall Air/Fuel Ratio 33.0 Molar throughout.

1 2 3 4 5 6 7 8 Line 1st. Stage 1st. Stage 1st. Stage Rich Gas Second Temp.after Therm.

No. Air/Fuel Combust. Pressure Temp.after Stage buming Effic.

Ratio.Mole Temp. Atmos. Expansion Pressure Rich Gas. Pct.

per Mole Deg. K Deg. K Atmos. Deg. K 1 33 1106 6.8 NA NA NA 31.48 2 3.13 1750 34.72 1272 6.8 1057 34.75 3 3.13 1850 56.87 1232 6.8 1051 36.16

Claims (4)

1. CLAIMS 1. A two stage gas turbine engine comprising a very hot first stage operating at higher pressure, in which fuel is bumed with air to give a gas rich in carbon monoxide and hydrogen, the gas is expanded through a turbine to lower pressure to give work and then bumed to completion in excess air and expanded again to give further work.The high content of carbon monoxide and hydrogen in the first stage makes possible the use of materials of turbine and casing construction there which would not resist the oxidative conditions at such high temperatures in conventional gas turbine operation.
2. 2. The materials of turbine blade and casing construction used in the hot first stage described in Claim 1 comprise high temperature metals and alloys or carbides, which would oxidise in the hot gas containing excess air in conventional engines.
3. 3. The materials in Claim 2 employed in construction of the first stage described in Claim 1 are specifically rhenium, tungsten and molybdenurn and their alloys with each other and other elements.
4. 4. The materials employed in construction of the first stage described in Claim 1 are also specifically the carbide of silicon and the carbides of the metals of Claim 3.

   4. The materials employed in construction of the first stage described in Claim 1 are also specifically the carbide of silicon and the carbides of the metals of Claim 3.

   Amendments to the claims have been filed as follows 1. A gas turbine engine in which a very high temperature first stage uses materials of construction for blades and casing which would be unstable in the oxidative! corrosive conditions which normally apply. The materials used will retain adequate strength and will not melt at the very high temperature. To achieve this the fuel is bumt initially in a deficiency of air as in a gasifier to yield a gas high in hydrogen and carbon monoxide: the required temperature and gas composition is achieved by control of airifuel ratio and air inlet pressure.After expansion of this very high temperature gas, more compressed air at lower pressure is added, combustion to completion in a second combustor at lower temperature takes place. and the gas is further expanded through a second turbine made of normal constructional metals. Efficiency is improved compared with normal practice. The invention applies to turbo-jet engines and non-regenerative gas turbines.

   2. The materials of turbine blade and casing construction used in the hot first stage described in Claim 1 comprise high melting temperature metals and and alloys or carbides, which would oxidise in the hot gas containing excess air in conventional engines.

   3. The materials in Claim 2 employed in construction of the first stage described in Claim 1 are specifically rhenium, tungsten and molybdenum and their alloys with each other and other elements.