US2243467A - Process and equipment for gas turbines - Google Patents

Process and equipment for gas turbines Download PDF

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US2243467A
US2243467A US158334A US15833437A US2243467A US 2243467 A US2243467 A US 2243467A US 158334 A US158334 A US 158334A US 15833437 A US15833437 A US 15833437A US 2243467 A US2243467 A US 2243467A
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turbine
combustion
compressor
fuel
air
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Jendrassik George
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C6/00Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas- turbine plants for special use
    • F02C6/003Gas-turbine plants with heaters between turbine stages

Description

May 27 1941.

G. JENDRASSIK PROCESS ANDEQUIPMENT FOR GAS TURBINES 6 Sheets-Sheet l I Filed Aug. 10, 1937 v invenfo a WW W'fnesses:

1&2.

May2z1941. GEM/$81K 2 243,467.

PROCESS AND EQUIPMENT FOR GAS TURBINES Filed Aug. 10, 1937 s Sheets-Sheet s a w 5 s2 u") w a" a u 1 0:

:9 Q -1- 8 Q 0% a Q y 1941- I G. JENDRASSIK Q I 2,243,467

PROCESS ANDYEQUIPMENT FOR GAS TURBINES Filed Aug. 10, 1937 6 Sheets-Sheet 4 Patented May 27, 1941 UNITED STATES PATENT orricr.

' 2,243,467 PROCESS ANDEQUIPLIENT FOB GAS George Jen'drassik, Budapest, Hungary Application'August 10, 1937, Serial No. 158,334 In Hungary. February 13, 1931 14 Claims.

In the case at gas turbines, in the first-place in the case of gas turbines operating with constant pressure, the employment, for the purpose of raising the efiiciency of the gas turbine, of heat exchange devices, constitutes 'a-known method. The heat exchange device required for this purpose represents, however, a structure of such large weight, whereby its employment for many purposesespecially in case a low total weight of structure is required-is rendered more diflicult. With the aid of the working process'according to the invention it is, even without any extensive employment of heat exchange devices, possiblewhilst keeping weights low-to obtain total 'eiilciencies of -45% and high outputs with as small dimensions as possible.

The gas turbine equipment necessary for this purpose, 1. e. for the realization of the new process is composed of a compressor, of a turbine and of a firing chamber or combustion chamber arranged I behind the compressor. According to the invention, in the compressor of the gas turbine the air is compressed to a certain advantageously chosen figure of pressure, following which combustion at constant pressureis produced so as to concentrate this combustion to a portion of. the compressed'air only or, heatis introduced into this portion or the air, which introduction of heatis I followed, after having led this portion of the working medium together with the other portion thereof in a common gas current into the turbine, by a further heat introduction taking place in the first stages of the turbine, during expansion, advantageously in such a manner that this first part of the expansion should be approxi- -mately isothermic, whilst the expansion in the remaining stages of the turbine shall take a course which approaches more the adiabatic than the isothermic, any introduction of heat during the expansion in the said remaining stages of the turbine being preferably dispensed with.

By selecting the pressures, i. e. the pressure drops of th various sections of expansion in a suitable manner it is possible to obtain the high efilciency and relatively high specific output men- 'tioned above, without having to raise the maxi-' mum mean temperature set up in 'the turbine above the limit to which materials of construction may still-safely be heated, i. e. above about 800 to 700 C. This is possible as'the said introduction of heat before the entranceinto the turrespectively.

Figs. 5-5:: and 6, respectively, are sections of two embodiments, shown by way. of example, of the gas turbine suitable for. carrying the working process into effect; Fig. 511 being a section along the section plane ia-ia through the com-. bustion chamber of Fig. 5,

Fig. '7 is a-longitudinal section of an embodiment, shown by way of example, of an atomizer capable of being employed advantageously for the suitable control of the gas turbine; finally,

blue takes place in a portion of the compressed working medium-only, at a temperature higher than the average. temperature of the entrance into the turbine, this being necessary in view of I Figs. 8 and 9 are diagrammatical longitudinal sections of such types of design of the gas turbine in which a separate low-pressure turbine behind the turbine according to the invention is I I also provided.

In the diagram according to Fig. 1 representing a mode, by way of example, of carrying the process according to the invention into effect, the

pressures set upin .the working process are traced -on the ordinate axis and the specific volumes of the gas on the abscissa axis-pa, 1m and To are denoting the initial pressure, specific volume and absolute temperature, respectively, of the gas drawn in by the compressor; m, m and T1 are the corresponding figuresrelating to the con-- dition of the gas when leaving the compressor, v: and Ta, this latter as a mean (equalization) temperature are the condition indicators of the gas after preliminary introduction of heat, whilst pa, in, T2 and pa, in, Ta are the condition indicators of the final condition of the expansion taking. place in the turbine during'further heat introduction at practically constant temperature, and those of the condition at the discharge from the turbine after the (adiabatic) 1 expansiomrespectively. in and D0 are, in general, equal, but

in cast throttling is employed in front of the compressor, Po will be lower than 9:.

The advantages f this Working process as compared to the workingprocesses known up to now may be seen from the Figs. 2, 3, and containing two sets oi curves. On these figures the ordinate is the compression ratio pi/po, whilst the abscissa is the proportion in which the pressure drop (pa-p) corresponding to nearly adiabatic expansion stands to the total pressure drop (pi-pa). The abscissa represents accordingly the so-called relative magnitude of the adiabatic expansion. Accordingly, on the first of the two sets of curves the ordinates belonging to the value zero of the abscissa represent the figures corresponding to the isothermic expansion, whilst the ordinates belonging to the value 1 of the abscissa represent the figures -belonging to the expansion without further introduction of heat. The curves traced in full lines belonging to one set of curves of the diagram are mutually connecting the points corresponding to an identical thermal efficiency, whilst the curves traced in broken lines, belonging to the other set of curves, are mutually connecting the points corresponding to certain constant figures of useful output of work (specific output of work) expressed in calories per kilogram of air drawn in. With starting data differing from the figures of temperature and of efliciency assumed (see Figs. 3 and 4 as compared with Fig. 2 the curves are of slightly different shape, but in other respect of generally similar character.

It appears from the diagram of Fig. 2 belonging to a compressor efliciency of 0.85 and a turbine eificiency of 0.9, further to a maximum temperature of 600 .C., that with a certain advantageous compression ratio, e. g. between p1/po=5 and pi/pn=11, a thermal efiiciency slightly more advantageous than 25 may be obtained also by meansof a process of work operating with purely adiabatic expansion, but the specific performance will only amount to a figure between 24 and 1'? calories per kg. If, as against this, after a section of expansion of a certain length at constant temperature, the introduction of heat is,

discontinued and the expansion is continued adiabatically, i. e. if the length of adiabatic expansion is shortened as compared to the known process, substantially mor advantageous thermal efficiencies and specific outputs are obtained. Thus for instance if the pressure ratio of the compression pilpo amounts to 15, whilst the relative length of adiabatic compression is 0.17, the thermal efiiciency will amount of 30.5% and the specific work to 40 cal./kg. By further diminishing the relative length of the adiabatic expansion, whilst keeping the ratio of compression pi/po constant, the efficiency will become lowered, but the specific output becomes improved still further. In view of the fact that the increase of the specific output is accompanied by a diminution of the own weight-of the mechanism, it is from this point of view advantageous to employ a relatively short adiabatic expansion.

With the conditions illustrated on Figs. 3 and 4, belonging both to a compressor efficiency of 0.9 and a turbine efliciency of 0.95, further to a maximum temperature of 600 C. and 700 0., respectively, the above considerations are, although with slightly different numerical values, valid likewise.

The smallest permissible length. of the adiabatic expansion will be determined by the consideration that the combustion should certainly, become completed already before discharge from the turbine, because any part of the fuel the combustion of which is effected after it has left the turbine represents a loss. Thus it will be preferable to employ an adiabatic expansion the rela-. tive length of which amounts to at least 0.05. The increase of the length of the adiabatic expansionwhilst keeping the ratio of compression constant-will be beyond a certain limit cause a diminution in efficiency as well as in specific performance of work; the specific performance will, however, diminish at a more rapid rate than the efficiency. The permissible maximum of the relative length of the adiabatic expansion can in the case of a greater ratio of compression not be as high in the case of a smaller ratio of compression because in this case the specific performance will become diminished more rapidly with the increase of the figures P2 Po lh-Po so as to be smaller than the figure resulting from the formula 92 Pi/Pu This result is obtained in the way that a boundary line, represented analytically by this formula, may be traced in the diagrams (as shown e. g. in Fig. 4 by the line a-b) beyond which line the conditions are unsatisfactory as regards the points of view explained above.

It also appears clearly from the curves of Fig. 2 that in case of the employment of adiabatic expansion of a magnitude of about 4/10, the efficiency is nearly independent of the length of the adiabatic expansion. For this reason it is advisable to proceed in such a manner when regulating the output of the gas turbine that whilst keeping the maximum temperature at a nearly invariable figure, the quantity of heat introduced is regulated by altering the magnitude of the adiabatic part. ceed, differently herefrom, in such a manner, that at the regulation the maximum temperature, i. e. the specific quantity of heat introduced at constant pressure is also varied. If it is desired to avoid that the temperatureshould become excessively diminished with the diminution of the output, it is also possible to proceed in such a manner, as to diminish the quantity of air supplied by the compressor by means of throttling, simultaneously with diminishing the quantity of heat introduced, which can be don by means of a simple throttling member employed in front of the compressor.

In Fig. 5 an embodiment of the gas turbine suitable for putting the working process according to the invention into effect is shown by way of example. of example on this drawing the compressor rotor 3 carrying-the rotor blades 2 is keyed jointly with the turbine rotor Scarrylng the turbine blades 4 It is, however, also possible to pro- In the embodiment shown by way I 2,243,467 on the turbine shaft 6 supported in the bearings 25, 26, both rotors being arranged inside the compressor and turbine casing l. Stationary blade rings I arelocated between successive blade rings of the compressor rotor, whilst stationary turbine blade rings 8 are located between the successive rotor blade rings of the turbine. The combustion space 9 into which the burners or atomizers lQ-lb' are opening, is arranged behind the compressor.

introduction of fuel are likewise arranged between the successiv blading stages of the turbine. The combustion space 9, the transverse section of which is shown on Fig. a,'is in the embodiment shown by way of example constituted by the interior of an annular hollow body (combustion chamber) 13 having a meridian section as shown on the figure, which is mounted in the turbine casing coaxially with the turbine axis, and the walls of which are-apart from a number of the turbine in the direction of the arrow l9 through the discharge opening 20. The purpose and operation of the combustion chamber l3 shown on Fig. 5 will be the'following: the air leaving the compressor is flowing rapidly and therefore, unless a furnace space of sufilcient magnitude is provided between the compressor and the turbine, insumcient time would be avail: able for the combustion of the fuel at the de- \sirable rate. on the other hand, it is all the more Further, the burners or atomizers I I, I2 for the necessary to make provision for the rapid combustion of the fuel because it is not in each case places of support-not entirely supported on the wall of the gas turbine casing, a clearance or channel l4 being left between them and the said gas turbine casing wall. The interior of the combustion chamber is fitted with openin s communicating with the space l5 leading from the compressor into the turbine and is fitted with the guiding or deflecting members IB'which stand opposite to the current of gas coming from the compressor. These guiding members may be shaped 'in the manner of a system of teeth preferably bent against the direction of flow of the gas current, and distributed over the whole periphery or located on certain parts of the periphery only, but they may possibly also be dispensed with entirely.

The feeding of the fuel injection members l0,

l0, II, I! etc. is effected e. g. by the reciprocating pump 12. If it is desired that the feeding should not be periodical, the balancing chests 23 e. g.

air chests per se known and commonly used in connection with reciprocating pumps may be inserted between the pump and the fuel injection members. The connection between the pump and the fuel injection members is effected by the ducts 24. The number and arrangement of the burners or atomizers employed for the p p se of introduction of heat are shown on the drawing in a diagrammatical manner only and may also be diflerent from those shown. The manner of operation of this apparatus is the following: In case of the rotation of the shaft 6 the compressor will draw in .air in the direction of the arrow ll through the inlet openings l8 and will compress this air by making it flow through the blade rings.

The members 10, In for the introduction of fuel are either leading hot gases of combustion into the combustion space 9 or directly passing the fuel into the said space; in the case of liquid fuel they are atomizing the latter; in the case of solid fuel it is not by means of atomizing but by some other means that the said members allow the fuel to pass into the space 9, whilst in the case of gaseous fuel the fuel, possibly already previously mixed with air will simply flow in through these members. The same is true also for the burners, or, as the case may be, atomizers H and I! for the introduction of fuel. Owing to the introduction of heat by the members l0, lo the temperature of the compressed air will become increased at constant pressure during the combustion, following which inflow into the turbine part will take place. In the turbine stages the gas will gradually expand and will be discharged from that the temperature of the air discharged from the compressor is high; in the case of a compression to ten times the original pressure apparatus according to the invention in which the guiding members l6 are wanting, the current of air will set the gas contained in the combustion chamber into rotation owing to friction whereby a layer of the air current adjacent to the rotating gas mixture will be detached and carried away into the combustion chamber, so that the entrance of fresh air into the combustion chamber will be assured in this case also, in consequence of which a corresponding quantity of' gas mixture will be displaced herefrom towards the entrance into the turbine. The same result is obtained and increased by the employment of t the guiding members I6, which are at certain places deflecting a part of the air current into the combustion chamber where the fresh air entering the chamber maintains a turbulent circulating motion. By selecting the magnitude of the communication opening in a suitable manner, and/or by giving a suitable shape to the guiding members it is possible to control the portion of the air supplied by the compressor which should enter the combustion chamber. The fuel is led into the combustion chamber by means of the members l0, l0 which are shaped in such a manner as to ensure the proper mixing of the fuel, this being done in the case of liquid fuel by means of atomizers; in the said combustion a degree as required according to what has been explained above. In view of the fact that it is only in a portion of the total quantity of air entering the combustion chamber that combustion takes place, the temperature ruling in the combustion chamber will be substantially higher than the average temperature of inlet into the turbine, and thus it will be possible to make satisfactory provision for self-ignition and for ensuring that the combustion should take place in a suitable degree before entrance into the turbine without raising the average temperature above the limit permissible in view of the heat resistance of the turbine structure. Thus if, for instance, the temperature of the gas leaving the compressor is 300 C., whilst that of the gas entering the turbine amounts, on the average, to 600 C., the. rise of temperature during the combustion taking place at constant pressure will In order to ensure that the walls of the com:

bustion chamber should not become excessively heated or should not heat the wall of the casing excessively, it will be advisable to allow a part of the air to flow through the clearances or ducts H left between the combustion chamber and the casing. The guidance of the air into this clearance is effected by means of the guiding member 2| which projects in the desired extent into the stream of air and directs a flow of suitable intensity towards the clearance. It is, however, also possible to insulate the wall of the combustion space by other means.

From the combustion chamber a quantity of hot gas corresponding to the quantity of air flowing in will flow out, which quantity of hot gas will accordingly become mixed with the rest of the current of air already before entering the turbine, so as to prevent that the turbine blades should become excessively heated at single places. Provision for the thorough mixing of the hot gases discharged from the chamber with the rest of the air should be made by leaving free a suitable mixing space as shown at 65 in Fig. 5.

As appears from the above it is a circumstance of great importance from the point of view of safe ignition and combustion, and on the other hand from the point of view of the proper moderation of the initial temperature of entrance into the turbine, that the current of air issuing from the compressor and entering the combustion chamber is divided into two parts, the combustion taking place first in one of these parts, whilst the other part serves for the reason referred to for the purposes of the admixture after combustion.

The fuel-introducing devices H and 12 which may of course be arranged, not only in the second and third stage but in any desired number and at any desired places are serving for the supplementary introduction of fuel for the purisothermic combustion into efiect the mannerin.

which the fuel is passed into the combustion space possesses extraordinary importance. How much fuel will be burned before entering the turbine and how much will be burned before entering the turbine and how much will be burnt in the turbine itself; is a matter which can be controlled by means of the perfection of the mixture, in the case of liquid material by the extent of atomization as well as by the selection of the particular place of the space of combustion at which the fuel is passed into the air, and furthermore also by the shape and/or dimensions given to the space of combustion. All effects by which pansion takes place.

combustion is slowed down, will also extend the length of the section along which isothermic ex- Thus it will be advisable to select and/or feed the atomizers, including the atomizers feeding into the combustion chamber, in such a manner, as to ensure that the degree of fineness of their atomization should be different. Fuel atomized into line particles will become burnt more quickly than fuel atomized into rough particles only and according to the degree of fineness of the atomization the combustion will extend into the turbine in a greater or in a less degree. At a certain degree of atomization, it is moreover possible to influence the progress of the combustion also by means of the shape given to the combustion chamber 9, and/or by means of the quantity of the fresh air introduced into the said chamber.

Should in the regulation of the gas-turbine the length of the isothermic section also be varied, it is preferable to vary at least one among the factors mentioned above. It is for this purpose that the constructional detail shown on Fig. 7, which represents an atomizer, serves. In the atomizer body 33 the valve 34 sliding in a tight manner preventing leakage is arranged, which valve by means of its conical end is regulating, according to the position occupied by the said valve, the cross section left open in the conical opening 36. It is through the bore-holes 3'! and 38 that the fuel enters the atomizer body from which it passes into the storage space- 39. In the storage space 39 the pressure of the fuel is such, that by acting on the lower surface of the valve 34 it will raise the said valve against the action of the spring 40. The preliminary tension of the spring 60 is controlled by adjusting the screwed spring support 4| in the longitudinal direction by turning thehandle 42., As the degree of atomization of the fuel depends on the pressure of the atomization, the amountof atomizations can also be controlled by turning the handle 42.

If the fuel is injected into the chamber of combustion by means of a plurality of atomizers, it will be preferable to effect regulation in case of the diminution of the output in such a. manner, that it is the feeding of the atomizers operating with a more rough atomization which is diminished first; whilst if in addition to the atomizers or fuel introducing burners feeding into the chamber of combustion fuel is introduced also a by means of burners'or atomizers arranged between the various stages of the turbine, it is in the first place the feeding of these last named burners or atomizers which is reduced in case of a diminution of the output.

' In case during the course of the regulation of 'the gas turbine the air drawn in is throttled, it

is necessary to arrange the throttling members,

e. g. the throttle-valve I4 (Fig. 6) in the inlet opening of the compressor.

The axial throughflow compressor and axial throughfiow turbine shownon Fig. 5 are particularly advantageousv for carrying the working process forming the subject of the invention into effect, because they permit the throughflow of a very large quantity of air with small dimensions of the apparatus, such small dimensions meaning at the sametime'a small weight of apparatus. This apparatus will, particularly in the caseof such a design in which it is substantially only the components ohperlpheral direction of the gas velocity which suffer any alteration, (in which accordingly, no alteration is suffered owing to the changes of energy .by the axial component) permit very high throughfiow speeds without substantial losses deriving therefrom.

It is characteristic for a compressor and/or turbine of such a type of design that the mean blade circle diameter of a stationary blade ring arranged between two adjacent rotating blade rings is approximately or exactly similar to the mean .figure of the blade circle diameter of the rotating blade rings. It is, moreover, also characteristic that the individual blade rings are arranged immediately alongside each other without any partition wall being provided between them.

In connection with the compressor it is," in order to obtain a particularly high efiiciency and high output, preferable to make provision for removing the spent boundary layer braked owing to friction and to rise of pressure from the surface of the casing and/or of the rotor, so as to. avoid the expected rise of pressure being prevented by the said layer. It is for this purpose that the arrangement per se known, shown by way of example in Fig. 5, serves, in which the spent boundary layer is through the ports 28, 28',

provided along the partition walls 21, 21' in some stage of the compressor, flowing back through the ducts 29, 29' and through ports 30, 30 to a place of lower pressure of the compressor, enter-.

ing into which place it will possess the regular contents of energy relatively to the conditions ruling there, and thus it will not be able to hinder that the rise of pressure should take place.

In view of the fact that the higher the temperature at which the turbine is operating, the higher will be its efiiciency, and the greater the peripheral velocity of its rotation, the smaller will be the dimensions of the turbine, the mechanical stresses to which the material of the rotor is subjected will also be substantial. For this reason it will be preferable to cool the rotorfrom inside. This can be carried into eifect in a suitable manner, if a current of air is allowed to. enter the that side of the combustion chamber 63 which faces the compressor there is provided opposite to the current of air leaving the compressor, and in such a manner as to project into this current of air, the guiding edge 10, which is able to divide the current .of .air leaving the compressor into two portions. The combustion chamber is fixed into the external casing H by meansof the guiding members I2, which guiding members-as already described in connection with Fig. 5-are arranged on various parts of the periphery of the combustion chamber and are intended for guiding the air-into the combustion chamber.

On the side facing the turbine of the combustion chamber body the ring rib I3 is closely approaching the first disc of the turbine rotor, so that only a small clearance remains between the two. The operation of this mechanism is the following: The current of air issuing from the compressor will, in'accordance with the arrows marked on the drawing, after striking against the guiding edge I0, fiow partly towards the combustionchamber, and partly flow inward in a radial direction alongside the combustionchamher through the duct 55. This part of the air current isthe coolingcurrent which will enter 'the turbine rotor in the axial direction through the openings 61, 61' provided in the discs, but will flow outwards along the turbine discs through the clearances or o enings left between the ring ribs 68, 68 of the iscs, and will thereby efiiciently cool the external surface of the said discs. The current of air returns into the space remaining between the turbine rotor and the chamber of combustion through the return bore-holes 69, 69 following which it will through interior 'of the rotor, and it is this purpose which is served by the openings, shown on'Fig. 5, on

the one hand in the rotor at 3|3l and on the other hand in the bearing pedestals of the casing at 32, 32'. a

In addition hereto it is possible to cool the turbine rotor also by means of the compressed air leaving the compressor, or by means of a portion of such air, the said portion, in this case, participating in the working cycle of the turbine after having passed through the interior of the turbine rotor. Moreover the external casing of the turbine also can be cooled by means of-the compressed air.

This last-named arrangement of the cooling of the turbine rotor'is shown on Fig. 6, according to which it is between the compressor rotor 6| ings, on a larger diameter, theribs 68, 68'

are provided. The annular ribs of the mutually adjacent discs are either not in mutual contact at all, or are in mutual contact in places only, so that a throughfiow cross-section-remains for the air. Beyond the annular ribs, on a diameter larger still, further througnfiow openings 69, 68' etc. are provided in the discs. On

the duct 66 flow into the turbine, mixed into the rest oithe working medium.

It is at the point of entrance into the com pressor,and in the low-pressure part of the turbine that the specific volume of the gas flowing through the gas turbine is the highest. In order to prevent that, particularly at the last-named place, an excessively long radial blade dimension should result, it is preferable to construct the turbine in such a manner that its low-pressure part should'beof greater diameter. 'In order tolprevent, on the other hand, that in this case its peripheral velocity should be too high, it is advisable to fix the speed of the low-pressure part at a lower figure. Such an arrangement is illustrated by Fig. 8. As shown in this figure, it isin the casing 63 that the rotor 44 of the turbine-compressor set is arranged, the said rotor being mounted so as to be rotatable in the bearings 4'|-41','on'the shaft 46 of the low pressure turbine rotor 45, which latter drives the enair propeller ergy consumer (e. g. the. airplane 51) in a direct manner.

The shaft 46 is supported by means of the bearings 48, 48, in' the casing 43. The air flows in at'the inlet opening 49 of the compressor, following which it flows through the compressor and through the high-pressure turbine 50 and proceeds into the low-pressure turbine 45 of greater diameter, A mechanical connection is effected between the rotors 44 and 45 bythe gear 5| in Fig. 9 the low-pressure rotor 52 of the com-j; pressor is likewise of greater diameter than the:

high-pressure rotor 53, and is keyed ,on a com-- mon shaft 56 with the compressor rotor 54 of the v turbine, as welhas with the consumer (air propeller) 51. Here the mechanical connection with the turbine-compressor set 53 is supplied by the gear wheel 55.

It is preferable to select the dimensions of the compressor-turbine set 44, or 53, respectively, and/or the distribution over high-pressure and low-pressure parts in such a manner that the resultant useful output of the part containing the high-pressure turbine (turbine and compressor should be zero, i. e. that the work of the turbine should just cover the requirements of energy of the compressor. In such a case it is preferable to take off the useful output at the shafts G or 56, respectively, and thus the output to be transmitted by the gears 55 or 5|, respectively, is likewise zero, or only very small. By driving the machines in such a manner the mechanical connection between the compressorturbine set and the low-pressure machines may also be dispensed with entirely, in which case the former rotates freely.

It is understood that the arrangements described and illustrated by the drawings. should only be considered to represent embodiments shown by way of example, innumerable other variants of arrangement being likewise suitable for carrying the process-forming the subject of the invention into effect, and it is only for the sake of brevity that such other variants of arrangement have not been described and illusing more the adiabatic than the isothermic,

ing fuel into the compressed working medium and burning a part thereof in a part of the working medium, i. e. so as to concentrate combustion to a portion of the working medium only, before the inlet into the turbine and by burning another part of the fuel during expan sion, in the gas current formed by the totality of the working medium, in the turbine itself, advantageously at a practically constant temperature, and in following this latter fuel combustion by a further expansion approaching more the adiabatic than the isothermic, preferably without any heat introduction.

2. A process for gas turbines, consisting in effecting the introduction of heat into the com-- pressed working medium by introducing and burning fuel in a combustion chamber, being arranged between the compressor and the turbine and suitable for holding only a part of the working medium, at constant pressure and thereafter by introducing and burning a further quantity of fuel in the gas current flowing in the turbine in a manner and distribution as, to ensure by its combustion advantageously approximately isothermic expansion, until an intermediate turbine stage, and in following this latter fuel combustion by an expansion, approaching more the adiabatic than the isothermic, preferably without any heat introduction.

3. A process for gas turbines, consisting in admitting the fuel and a part of the compressed combustion air entering the turbine to a cornbustion space of constant pressure between the compressor and the turbine in a manner and quantity ensuring the commencement of the 4. A process for gas turbines, consisting in effecting the introduction of heat by introducing fuel into the compressed working medium and burning a part thereof in a part of the working medium, i. to a port-ion of the working medium only, before the inlet into the turbine and by burning another part of the fuel during expansion, in the gas current formed by the totality of the working medium, in the turbine itself, advantageously at a practically constant temperature, and in following this latter fuel combustion by a further expansion approaching more the adiabatic than the isothermic, preferably without any heat introduction, the magnitude of the expansion without any heat introduction being chosen so as to keep the relative magnitude ("l-Po of this expansion (the proportion of the pressure drop pe -p0 thereof to the full pressure drop 191-770 in the turbine) between the limits 0.05 and 5. A process for gas turbines, consisting in effecting the introduction of heat by introducing fuel into the compressed working medium and burning a part thereof in a part of the working medium, i. e. so as to concentrate combustion to a portion of the working medium only, before the inlet into the turbine and by burning another part of the fuel during expansion, in the gas current formed by the totality of the working medium, in the turbine itself, advantageously at a practically constant temperature, in following this latter fuel combustion by a further expansion approaching more the adiabatic than the isothermic, preferably without any heat introduction, and in altering the relative magnitude 2 1-7 0 of this further expansion (the proportion of the pressure drop pzpn thereof to the full pressure drop pipo in the turbine) with the alteration of the heat quantity introduced when the output of the turbine is altered.

bine stage, advantageously at approximately isothermic expansion, this section of expansion being followed by a further expansionapproach- "following this latter fuel combustion by a further expansion approaching more the adiabatic than the isothermic, and in throttling the air to be compressed simultaneously with the diminution of the heat quantity introduced when the output of the turbine is diminished.

7. A process for gas turbines, consisting in admitting the liquid fuel and a part of the compressed combustion air enteringthe turbine to a combustion space of constant pressure between the compressor and the turbine in a manner and quantity ensuring the commencement of t e combustion in the said combustion chamc. so as to concentrate combustion 1 mjhltbine, an annular hollowbody. around theturbine shaft forming combustion chamber the-said compressor and turbine, suither .and the continuation thereof in the gas.

current formed by the totality of the working medium, in the turbine itself, until an intermediate turbine stage, advantageously at approximately isothermic expansion, in following this expansion with a further expansion without any heat introduction, and in altering the degree of atomization of the fuel in the. combustion air when the output of the turbine is altered.

8. A process for gas turbines, consisting in ad mitting the liquid fuel, by means ofatomizers;

the said hollow body close to the working medium current flowing from the compressor towards the turbine, a by-pass air gap between the said hollow body and the engine casing, wall prolonga-' tions projecting into the compressed air flow and deflecting a part thereof into the said gap, provided on the edge of the inlet opening of the combustion chamber nearer the compressor,

, means suitable to deflect at least a part of the and a part of the compressed combustion air entering the turbine to a combustion space of constant pressure between the compressor and the turbine in a manner and quantity ensuring ,the commencement of the combustion in the said combustion chamber and the continuation thereof in the gas current formed by the totality of the working medium, in the turbine itself, until an remaining quantity of the working medium current. into thesaid combustion chamber, situated on the edge of the said annular inlet opening nearer the turbine, and means supplying fuel into the working medium at best until an intermediate stage of the turbine ina manner and ditsribution ensuring-"combustion partly .before the inflow into the turbine in the said combustion chamber, and partly during the first part of the expansion in the turbine also, this first intermediate turbine stage, advantageously at apf proximately isothermicexpansion, and in diminishing firstly the fuel supply of the atomizers supplying between the turbine stages when the output of the turbine is diminished, the said sec-' tion of expansion being followed by a fiu'ther expansion approaching more the adiabatic than the isothermic, preferably without any heat introduction.

9. Inan equipment for gas turbines, a compressor for compressing the working medium, a

. tageously at isothermic expansion, and accompanied by a further expansion approaching more the adiabatic than-the isothermic, prefe r ably without any heat introduction.

10. In an equipment for gas turbines, a compressor for compressing the working medium, a gas turbine, an annular hollow body around the mane shaft forming the'combustion chamber' between, the said compressor and turbine, suitable? for holding only a part of the compressed part of expansion being effected advantageously isothermically; I

- 12. In an equipment for'g'as turbines, a compressor for compressing the working medium, a gas turbine fed with liquidfuel, a combustion chamber joining on to theflow path of the working medium between the said compressor and turbine suitable for holding only a part of the compressed working medium' streaming to the v inlet into the turbine, and atomizers of different de ree of atomization for supplying. the'fuel into thesaid combustion chamber and into the first stages of the turbine in a manner ensuring com bustion and approximately isothermic expansion until an intermediate expansion stage.

13. In an equipment for gas turbines, a compressor for compressing the working medium having blade rows of average diameters equal to the mean figures of the average diameters of the adjacent blade rows, a gas turbine, a combustion chamber joining on to the flow path of the working medium between the said compressorand turbine suitable for holding only a part of the compressed working medium streaming to the inlet into the turbine, and means introducing fuel into the compressed working medium in a manner ensuring combustion to be started before the inflow into the turbine in the said combustion chamber and completed in an intermediate expansion stage of the turbine, the combustion workingfimedium, an annularinlet opening-on the said hollow-body close to the current of working medium flowing from the compressor to-- wards the turbine, means suitable to deflect at least a part of the working medium current into the combustion chamber, situated on the edge of the saidannular inlet opening facing'the flow and'in'eans supplying fuel at least partlyinto the said combustion chamberin amanner and distribution ensuring combustion partlybefore. the the turbine, and partly in the first inflowinto v 7 stages of the-turbine also, this latter combustion being eflected advantageously at isothermic ex-. pansion: A

11-. In an equipment for gas turbines, a com-. for compressing the "work n medium, a

ing'the turbine being effected advantageously at isothermic expansion. i v

14. In an equipment for. gas turbines, a compressor for compressing the working medium having blade rows of average diameters equal to ableforholding-lonly apart of the compressed an annular inlet opening on the mean'figures of the average diameters of. the adjacent blade rows and channels in the walls of its working space connecting zones of lower and higherpressure, a gas turbine, a combustion chamber joining on to the. flow path of the workv ing medium between the said compressor and turbine suitable for holding only a part of the 1 compressed'working medium streaming to the inlet into the turbine, and means introducing fuel into the compressed worklng medium in a manner ensuring combustion to be started before the inflow. into the turbine in the said combustion chamber and completed in an intermediate .ex-

.pansion stage of the turbine, the combustion in the turbine being effected advantageously. at

isothermic expansion

US158334A 1937-02-13 1937-08-10 Process and equipment for gas turbines Expired - Lifetime US2243467A (en)

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Cited By (62)

* Cited by examiner, † Cited by third party
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US2422213A (en) * 1944-06-09 1947-06-17 Westinghouse Electric Corp Combustion chamber
US2454738A (en) * 1944-01-31 1948-11-23 Power Jets Res And Development Internal-combustion turbine power plant
US2468461A (en) * 1943-05-22 1949-04-26 Lockheed Aircraft Corp Nozzle ring construction for turbopower plants
US2474143A (en) * 1944-07-13 1949-06-21 Fairey Aviat Co Ltd Propulsion means for aircraft and the like
US2479143A (en) * 1944-12-07 1949-08-16 Jr Samuel W Traylor Gas turbine
US2479056A (en) * 1944-08-23 1949-08-16 United Aircraft Corp Cooling turbine rotors
US2479777A (en) * 1943-05-22 1949-08-23 Lockheed Aircraft Corp Fuel injection means for gas turbine power plants for aircraft
US2487514A (en) * 1943-01-16 1949-11-08 Jarvis C Marble Turbine rotor cooling
US2489683A (en) * 1943-11-19 1949-11-29 Edward A Stalker Turbine
US2520967A (en) * 1948-01-16 1950-09-05 Heinz E Schmitt Turbojet engine with afterburner and fuel control system therefor
US2526409A (en) * 1945-01-09 1950-10-17 Lockheed Aircraft Corp Turbo-propeller type power plant having radial flow exhaust turbine means
US2532721A (en) * 1944-08-23 1950-12-05 United Aircraft Corp Cooling turbine rotor
US2548804A (en) * 1945-03-23 1951-04-10 Stewart Warner Corp Jet propulsion apparatus
US2552239A (en) * 1946-10-29 1951-05-08 Gen Electric Turbine rotor cooling arrangement
US2555924A (en) * 1948-11-27 1951-06-05 Bbc Brown Boveri & Cie Fluid cooled rotor structure
US2563744A (en) * 1942-03-06 1951-08-07 Lockheed Aircraft Corp Gas turbine power plant having internal cooling means
US2578481A (en) * 1946-03-25 1951-12-11 Rolls Royce Gas turbine power plant with auxiliary compressor supplying cooling air for the turbine
US2579049A (en) * 1949-02-04 1951-12-18 Nathan C Price Rotating combustion products generator and turbine of the continuous combustion type
US2583872A (en) * 1947-08-02 1952-01-29 United Aircraft Corp Gas turbine power plant, including planetary gearing between a compressor, turbine, and power consumer
US2587649A (en) * 1946-10-18 1952-03-04 Pope Francis Selective turbopropeller jet power plant for aircraft
US2589078A (en) * 1944-03-29 1952-03-11 Power Jets Res & Dev Ltd Aircraft propulsion power plant
US2589853A (en) * 1946-03-12 1952-03-18 Bristol Aeroplane Co Ltd Aircraft power plant having two or more gas turbine power units to drive one or more airscrews in various combinations
US2603453A (en) * 1946-09-11 1952-07-15 Curtiss Wright Corp Cooling means for turbines
US2613501A (en) * 1945-06-02 1952-10-14 Lockheed Aircraft Corp Internal-combustion turbine power plant
US2618461A (en) * 1948-10-05 1952-11-18 English Electric Co Ltd Gas turbine
US2619797A (en) * 1948-01-28 1952-12-02 Rolls Royce Gas turbine engine driving a propeller
US2620624A (en) * 1952-12-09 wislicenus
US2623357A (en) * 1945-09-06 1952-12-30 Birmann Rudolph Gas turbine power plant having means to cool and means to compress combustion products passing through the turbine
US2627717A (en) * 1948-06-11 1953-02-10 Laval Steam Turbine Co Multiple gas turbine power plant having speed governors to bypass power turbine and regulate fuel feed
US2630678A (en) * 1947-08-18 1953-03-10 United Aircraft Corp Gas turbine power plant with fuel injection between compressor stages
US2630676A (en) * 1947-01-20 1953-03-10 Donald W Seifert Axial flow jet motor with rotating combustion products generator and turbine
US2630677A (en) * 1947-01-20 1953-03-10 Donald W Seifert Axial flow jet motor with reversely rotating continuous combustion type combustion products generator and turbine
US2631429A (en) * 1948-06-08 1953-03-17 Jr Harold M Jacklin Cooling arrangement for radial flow gas turbines having coaxial combustors
US2638741A (en) * 1948-08-11 1953-05-19 Jr Henry M Putman Axial flow gas turbine having reheating means and specially shaped rotor and stator blades to provide isothermal expansion
US2640319A (en) * 1949-02-12 1953-06-02 Packard Motor Car Co Cooling of gas turbines
US2645412A (en) * 1949-09-28 1953-07-14 United Aircraft Corp Locking device for split-compressor type turbopower plants
US2648519A (en) * 1948-04-22 1953-08-11 Campini Secondo Cooling combustion turbines
US2653446A (en) * 1948-06-05 1953-09-29 Lockheed Aircraft Corp Compressor and fuel control system for high-pressure gas turbine power plants
US2654220A (en) * 1943-12-01 1953-10-06 Jarvis C Marble Apparatus for directing air to combustion products turbines
US2669420A (en) * 1948-07-03 1954-02-16 Kellogg M W Co Turbine structure
US2675195A (en) * 1949-12-14 1954-04-13 Gerard P Herrick Power plant for aircraft having vertical as well as horizontal propulsion
US2677932A (en) * 1948-08-27 1954-05-11 Gen Electric Combustion power plants in parallel
US2689681A (en) * 1949-09-17 1954-09-21 United Aircraft Corp Reversely rotating screw type multiple impeller compressor
US2702665A (en) * 1951-03-07 1955-02-22 United Aircraft Corp Stator construction for axial flow compressors
US2786625A (en) * 1950-08-01 1957-03-26 Rolls Royce Turbo-machines
US2944397A (en) * 1951-03-23 1960-07-12 American Mach & Foundry Combustion chambers for gas turbine power plants
US2947143A (en) * 1952-10-15 1960-08-02 Nat Res Dev Baffle arrangement for combustion equipment
US2977760A (en) * 1955-03-16 1961-04-04 Bristol Aero Engines Ltd Annular combustion chambers for use with compressors capable of discharging combustion supporting medium with a rotary swirl through an annular outlet
US2989849A (en) * 1951-08-04 1961-06-27 United Aircraft Corp Fuel control system for a twin spool gas turbine power plant
US3055179A (en) * 1958-03-05 1962-09-25 Rolls Royce Gas turbine engine combustion equipment including multiple air inlets and fuel injection means
US3088281A (en) * 1956-04-03 1963-05-07 Bristol Siddeley Engines Ltd Combustion chambers for use with swirling combustion supporting medium
US3333414A (en) * 1965-10-13 1967-08-01 United Aircraft Canada Aerodynamic-flow reverser and smoother
US3940923A (en) * 1971-05-13 1976-03-02 Engelhard Minerals & Chemicals Corporation Method of operating catalytically supported thermal combustion system
US4197700A (en) * 1976-10-13 1980-04-15 Jahnig Charles E Gas turbine power system with fuel injection and combustion catalyst
WO1980001591A1 (en) * 1979-02-06 1980-08-07 C Jahnig Gas turbine power system with fuel injection and combustion catalyst
US4893468A (en) * 1987-11-30 1990-01-16 General Electric Company Emissions control for gas turbine engine
DE4236071A1 (en) * 1992-10-26 1994-04-28 Abb Research Ltd A method for a multi-stage combustion in gas turbines
US6260349B1 (en) 2000-03-17 2001-07-17 Kenneth F. Griffiths Multi-stage turbo-machines with specific blade dimension ratios
US6298653B1 (en) 1996-12-16 2001-10-09 Ramgen Power Systems, Inc. Ramjet engine for power generation
US6347507B1 (en) 1992-09-14 2002-02-19 Ramgen Power Systems, Inc. Method and apparatus for power generation using rotating ramjets
US6378287B2 (en) 2000-03-17 2002-04-30 Kenneth F. Griffiths Multi-stage turbomachine and design method
US6446425B1 (en) 1998-06-17 2002-09-10 Ramgen Power Systems, Inc. Ramjet engine for power generation

Cited By (65)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2620624A (en) * 1952-12-09 wislicenus
US2563744A (en) * 1942-03-06 1951-08-07 Lockheed Aircraft Corp Gas turbine power plant having internal cooling means
US2487514A (en) * 1943-01-16 1949-11-08 Jarvis C Marble Turbine rotor cooling
US2479777A (en) * 1943-05-22 1949-08-23 Lockheed Aircraft Corp Fuel injection means for gas turbine power plants for aircraft
US2468461A (en) * 1943-05-22 1949-04-26 Lockheed Aircraft Corp Nozzle ring construction for turbopower plants
US2489683A (en) * 1943-11-19 1949-11-29 Edward A Stalker Turbine
US2654220A (en) * 1943-12-01 1953-10-06 Jarvis C Marble Apparatus for directing air to combustion products turbines
US2454738A (en) * 1944-01-31 1948-11-23 Power Jets Res And Development Internal-combustion turbine power plant
US2589078A (en) * 1944-03-29 1952-03-11 Power Jets Res & Dev Ltd Aircraft propulsion power plant
US2422213A (en) * 1944-06-09 1947-06-17 Westinghouse Electric Corp Combustion chamber
US2474143A (en) * 1944-07-13 1949-06-21 Fairey Aviat Co Ltd Propulsion means for aircraft and the like
US2479056A (en) * 1944-08-23 1949-08-16 United Aircraft Corp Cooling turbine rotors
US2532721A (en) * 1944-08-23 1950-12-05 United Aircraft Corp Cooling turbine rotor
US2479143A (en) * 1944-12-07 1949-08-16 Jr Samuel W Traylor Gas turbine
US2526409A (en) * 1945-01-09 1950-10-17 Lockheed Aircraft Corp Turbo-propeller type power plant having radial flow exhaust turbine means
US2548804A (en) * 1945-03-23 1951-04-10 Stewart Warner Corp Jet propulsion apparatus
US2613501A (en) * 1945-06-02 1952-10-14 Lockheed Aircraft Corp Internal-combustion turbine power plant
US2623357A (en) * 1945-09-06 1952-12-30 Birmann Rudolph Gas turbine power plant having means to cool and means to compress combustion products passing through the turbine
US2589853A (en) * 1946-03-12 1952-03-18 Bristol Aeroplane Co Ltd Aircraft power plant having two or more gas turbine power units to drive one or more airscrews in various combinations
US2578481A (en) * 1946-03-25 1951-12-11 Rolls Royce Gas turbine power plant with auxiliary compressor supplying cooling air for the turbine
US2603453A (en) * 1946-09-11 1952-07-15 Curtiss Wright Corp Cooling means for turbines
US2587649A (en) * 1946-10-18 1952-03-04 Pope Francis Selective turbopropeller jet power plant for aircraft
US2552239A (en) * 1946-10-29 1951-05-08 Gen Electric Turbine rotor cooling arrangement
US2630676A (en) * 1947-01-20 1953-03-10 Donald W Seifert Axial flow jet motor with rotating combustion products generator and turbine
US2630677A (en) * 1947-01-20 1953-03-10 Donald W Seifert Axial flow jet motor with reversely rotating continuous combustion type combustion products generator and turbine
US2583872A (en) * 1947-08-02 1952-01-29 United Aircraft Corp Gas turbine power plant, including planetary gearing between a compressor, turbine, and power consumer
US2630678A (en) * 1947-08-18 1953-03-10 United Aircraft Corp Gas turbine power plant with fuel injection between compressor stages
US2520967A (en) * 1948-01-16 1950-09-05 Heinz E Schmitt Turbojet engine with afterburner and fuel control system therefor
US2619797A (en) * 1948-01-28 1952-12-02 Rolls Royce Gas turbine engine driving a propeller
US2648519A (en) * 1948-04-22 1953-08-11 Campini Secondo Cooling combustion turbines
US2653446A (en) * 1948-06-05 1953-09-29 Lockheed Aircraft Corp Compressor and fuel control system for high-pressure gas turbine power plants
US2631429A (en) * 1948-06-08 1953-03-17 Jr Harold M Jacklin Cooling arrangement for radial flow gas turbines having coaxial combustors
US2627717A (en) * 1948-06-11 1953-02-10 Laval Steam Turbine Co Multiple gas turbine power plant having speed governors to bypass power turbine and regulate fuel feed
US2669420A (en) * 1948-07-03 1954-02-16 Kellogg M W Co Turbine structure
US2638741A (en) * 1948-08-11 1953-05-19 Jr Henry M Putman Axial flow gas turbine having reheating means and specially shaped rotor and stator blades to provide isothermal expansion
US2677932A (en) * 1948-08-27 1954-05-11 Gen Electric Combustion power plants in parallel
US2618461A (en) * 1948-10-05 1952-11-18 English Electric Co Ltd Gas turbine
US2555924A (en) * 1948-11-27 1951-06-05 Bbc Brown Boveri & Cie Fluid cooled rotor structure
US2579049A (en) * 1949-02-04 1951-12-18 Nathan C Price Rotating combustion products generator and turbine of the continuous combustion type
US2640319A (en) * 1949-02-12 1953-06-02 Packard Motor Car Co Cooling of gas turbines
US2689681A (en) * 1949-09-17 1954-09-21 United Aircraft Corp Reversely rotating screw type multiple impeller compressor
US2645412A (en) * 1949-09-28 1953-07-14 United Aircraft Corp Locking device for split-compressor type turbopower plants
US2675195A (en) * 1949-12-14 1954-04-13 Gerard P Herrick Power plant for aircraft having vertical as well as horizontal propulsion
US2786625A (en) * 1950-08-01 1957-03-26 Rolls Royce Turbo-machines
US2702665A (en) * 1951-03-07 1955-02-22 United Aircraft Corp Stator construction for axial flow compressors
US2944397A (en) * 1951-03-23 1960-07-12 American Mach & Foundry Combustion chambers for gas turbine power plants
US2989849A (en) * 1951-08-04 1961-06-27 United Aircraft Corp Fuel control system for a twin spool gas turbine power plant
US2947143A (en) * 1952-10-15 1960-08-02 Nat Res Dev Baffle arrangement for combustion equipment
US2977760A (en) * 1955-03-16 1961-04-04 Bristol Aero Engines Ltd Annular combustion chambers for use with compressors capable of discharging combustion supporting medium with a rotary swirl through an annular outlet
US3088281A (en) * 1956-04-03 1963-05-07 Bristol Siddeley Engines Ltd Combustion chambers for use with swirling combustion supporting medium
US3055179A (en) * 1958-03-05 1962-09-25 Rolls Royce Gas turbine engine combustion equipment including multiple air inlets and fuel injection means
US3333414A (en) * 1965-10-13 1967-08-01 United Aircraft Canada Aerodynamic-flow reverser and smoother
US3940923A (en) * 1971-05-13 1976-03-02 Engelhard Minerals & Chemicals Corporation Method of operating catalytically supported thermal combustion system
US4197700A (en) * 1976-10-13 1980-04-15 Jahnig Charles E Gas turbine power system with fuel injection and combustion catalyst
WO1980001591A1 (en) * 1979-02-06 1980-08-07 C Jahnig Gas turbine power system with fuel injection and combustion catalyst
US4893468A (en) * 1987-11-30 1990-01-16 General Electric Company Emissions control for gas turbine engine
US6347507B1 (en) 1992-09-14 2002-02-19 Ramgen Power Systems, Inc. Method and apparatus for power generation using rotating ramjets
US6510683B1 (en) 1992-09-14 2003-01-28 Ramgen Power Systems, Inc. Apparatus for power generation with low drag rotor and ramjet assembly
DE4236071A1 (en) * 1992-10-26 1994-04-28 Abb Research Ltd A method for a multi-stage combustion in gas turbines
DE4236071C2 (en) * 1992-10-26 2002-12-12 Alstom A method for a multi-stage combustion in gas turbines
US6298653B1 (en) 1996-12-16 2001-10-09 Ramgen Power Systems, Inc. Ramjet engine for power generation
US6434924B1 (en) 1996-12-16 2002-08-20 Ramgen Power Systems, Inc. Ramjet engine for power generation
US6446425B1 (en) 1998-06-17 2002-09-10 Ramgen Power Systems, Inc. Ramjet engine for power generation
US6378287B2 (en) 2000-03-17 2002-04-30 Kenneth F. Griffiths Multi-stage turbomachine and design method
US6260349B1 (en) 2000-03-17 2001-07-17 Kenneth F. Griffiths Multi-stage turbo-machines with specific blade dimension ratios

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