The invention relates to a gas turbine having a casing surrounding an impeller with several compressor blades disposed in an entry area of the casing for

compressing air, at least one combustion chamber downstream the impeller for combusting a mixture of compressed air and injected fuel and several turbine blades downstream the at least one combustion chamber and disposed in an exit area of the casing. The invention furthermore relates to a method for operating a gas turbine.

The typical construction of a gas turbine and method of operating a gas turbine is given by compressing air entering a gas turbine by means of compressor blades rotating as an impeller in the entry area of a casing being stationary with respect to a frame of reference that is for example given by the earth or any kind of propelled vehicle. The compressed air is mixed with fuel injected into the stationary casing and combusted in the stationary combustion chamber downstream the impeller and situated in the casing. The combusted air-fuel-mixture is exhausted through turbine blades being positioned in the exit area of the casing and being rotated in/underneath the casing. The turbine blades are driving a shaft on which the impeller is positioned.

These kind of gas turbines of small to medium thrust class which are typically used for example as aircraft engines are limited in overall pressure ratio, as well as in bypass ratio due to their overall size. It is known that a high overall pressure ratio leads to a high thermal efficiency and a high bypass ratio leads to a high propulsive efficiency. The overall efficiency and thus the fuel consumption of such gas turbines are determined by these two efficiencies. With increasing bypass ratios, the jet noise is also reduced due to the speed reduction of the core and bypass jet.

These gas turbines of small to medium-thrust class do not achieve the high overall pressure ratios and bypass ratios as the engines of larger thrust class. This situation exists, because many of the geometrical and aerodynamic parameters of the compression and expansion systems of the turbo machine do not scale down proportionally with decreasing size. Tip clearance between impeller and casing as well as secondary losses at the endwalls are examples. Furthermore, adequate profile accuracy with small blades is difficult to achieve just as trailing edge thickness is limited down to some absolute value. Also turbine cooling channels become too small to be cast.

Especially the use of axial compressors in small gas turbines is limited by these effects. Consequently, in almost all small gas turbines centrifugal compressors are used, which also have relatively large tip clearance losses and therefore limited compressor efficiencies, but they can handle higher pressure ratios per stage. But for both, the axial as well as the radial configuration, the achievable compressor pressure ratio is limited for small gas turbines, whereby only small thermal efficiencies can be achieved.

In vehicle engines just like in stationary gas turbines for power plants, the pollutant emission and fuel consumption is of central importance due to environmental specifications and operating costs. Both can be reduced by increased efficiencies of the gas turbine. Besides the pressure ratio, the thermal efficiency of gas turbines is significantly limited by the achievable turbine inlet temperature, which is mainly limited by the possible thermal loading of the first turbine stage. Until now, the cooling technology of turbine stages was improved, in order to increase the turbine inlet temperature. In today's vehicle engines, in particular aircraft engines up to 20% of the air flowing through the core engine will be used for turbine cooling. This cooling air is considered thermodynamically as a loss, because this air flow is not participating in the combustion process and thereby is not being energized, while it still has to be pressurized in the compressor. The insertion of the cooling channels in small turbine geometries is problematic because of manufacturing reasons, whereby small gas turbines cannot use turbine cooling. Thus, the turbine inlet temperature of small gas turbines is limited by the tolerable temperature of the high-temperature Ni-based alloys, for instance.

In order to achieve similar overall pressure ratios and bypass ratios as engines of high-thrust class do, it is necessary to develop a core engine, which eliminates or reduces the main disadvantages of small gas turbine engines.

It is an object of the present invention to provide a novel gas turbine and a novel method of operating a gas turbine in order to overcome the mentioned

disadvantages of small gas turbine engines. In particular it is an object of the invention to provide a gas turbine having higher turbine inlet temperatures in combination with higher compressor and turbine efficiencies.

In a gas turbine according to the invention all the compressor blades and all the turbine blades are fixed at their respective outer radial ends on the inner wall of the casing, thus forming a unitary rotatable system comprising a rotatable combustion chamber, in particular by mounting the casing rotatably directly or indirectly at a fixation member by means of bearings. By means of such a fixation member the casing is rotatably mounted in any kind of frame of reference, for example mounted to the earth (in particular for power plants) or mounted in vehicles.

Using such a construction a method of operating a gas turbine may be achieved wherein air is entering a rotating casing, the air is compressed by means of compressor blades being fixed at their respective outer radial ends on the inner wall of the casing, the compressed air is entering at least one combustion chamber being rotated together with the casing and the compressor blades and the air being mixed with fuel and combusted and wherein the expanding combusted air/fuel mixture is exhausted through turbine blades being also fixed at their respective outer radial ends on the inner wall of the casing and being rotated together with the casing, the at least one combustion chamber and compressor blades.

The most essential feature of this invention is the use of a rotatable / rotating casing, thus the blades of the impeller and the turbine section can be attached to the inner wall of the rotatable casing with their respective outer radial ends. As a consequence all the blades are not excelled to tensile stress when rotated. This construction allows not only the use of a metallic material but also the use of ceramics for the compressor blades and or the turbine blades and/or the combustion chamber. In particular when using ceramics high turbine inlet temperatures in the absence of turbine cooling, plus after-burning within the (ceramic) turbine is possible.

The problem in using ceramic materials is that these materials can handle compressive stresses much better than tensile stress. In conventional turbine rotors, the turbine blades are attached to the hub and have to tolerate high tensile stresses due to the centrifugal force. In contrast to common gas turbines the turbine blades, compressor blades and combustion chamber walls (if there are any) in the invention are attached to the casing. The radial outward directed centrifugal force leads to compressive stress in the blade roots, which are mounted in the rotating casing.

In conventional gas turbines the centrifugal load is being held by the turbine and compressor discs. Since in this invention the compressor blade, the turbine blades and the combustion chamber walls are attached to the rotating casing, no discs may be used in the whole design. With this architecture, the casing is a particularly highly loaded component since it rotates and since it must hold back the centrifugal forces of the blades and combustor walls. If in specific constructions neither metallic high temperature materials nor high- temperature ceramics may handle the high mechanical stresses in the rotating casing then also the use of fiber composite materials may be used. Fibre composite materials CFRP (Carbon Fiber Reinforced Plastic) or CMC ceramics (Ceramic Matrix Composites) may be used, in particular for the casing walls.

CFRP tolerates only low component temperatures, whereby the practical usability of CFRP is limited. CMC ceramics are high-temperature ceramics, which are mechanically reinforced by ceramic fiber. Thus, the good mechanical properties of fiber composite material can be combined with the good high temperature properties of ceramic materials. Today, the manufacturing of complex geometries from CMC ceramics is still not possible; however, the production of tube

geometries is well engineered. This circumstance is beneficial for the invention, because a load-bearing tube of CMC ceramic can be placed around the rotating casing. The blades and blade connections can be made from conventional high- temperature materials, in particular ceramics.

If the geometry is made of metallic materials, it is well suited for a rapid

manufacturing scheme such as Selective Laser Melting (SLM), which enables integral geometries with a high degree of freedom.

The losses in turbomachinery are generally given by blade profile losses, tip clearance losses between stationary and rotating components and secondary flow losses on inner and outer endwalls.

Especially for small gas turbines of common construction the tip clearance losses, which arise between the stationary and rotating components, have a particularly large influence. The blades of the last compressor stage are very small. Because of manufacturing reasons the tip clearance cannot be reduced in this manner as it would be necessary to comply with the geometric similarity of large gas turbines. Hence, the tip clearance relative to the blade height is significantly larger, which drastically reduces the efficiency of small turbo machines of common construction. In the presented invention the blades will be attached to the casing avoiding any tip clearance and corresponding losses, as mentioned. Whereby any gaps between the outer radial ends of the blades as known from common gas turbines are omitted, providing less loss.

Analogously to common gas turbines, the gas turbine of the invention consists of a compressor, in particular a single stage centrifugal compressor, a combustion chamber and a turbine, in particular a single stage radial flow turbine. The specific feature of this gas turbine is the rotating combustion chamber. It is placed directly between the impeller of the centrifugal compressor and the rotor of the turbine. Thus, the combustion takes place in the rotating system. This furthermore allows a construction in which no stator vanes are required in the turbine and the compressor stage providing a major advantage of the rotating combustion chamber, because these components are thermally highly loaded and very costly in production and maintenance.

In the gas turbine of the invention kinetic energy at the exit of the compressor impeller does not have to be converted to static pressure. High absolute velocities occur at the inflow of the combustion chamber, while relative velocities are low enabling optimal conditions for a stable combustion.

In contrast to the invention in ordinary gas turbines a stator needs to be placed downstream of the compressor impeller to convert the high kinetic energy at the impeller outlet into static pressure at the stator outlet. In front of the turbine rotor a stator will be placed to induce a swirl component to the fluid, which can be extracted in the turbine rotor again.

By omitting the stator vanes in the invention friction losses can be avoided, because no profile losses will be induced at the stator vanes. Generally speaking, centrifugal compressors have a lower efficiency than axial compressors and the centrifugal compressor's stator produces the main loss. Centrifugal compressors typically achieve efficiencies up to 85%. However, the impeller itself can achieve efficiencies over 90%. This is a further advantage of the invention, since only the high efficiency impeller may be used. The largest sources of loss are eliminated due to the avoidance of turbine cooling, tip clearance and stator vanes.

Consequently, this invention is predestined for the use in small applications because most small gas turbines are largely affected by the described loss mechanisms.

In a preferred embodiment of the invention the compressor blades and turbine blades may extend from the inner wall of the casing towards the outer wall of a tubular shroud and may be fixed to it at their respective inner radial ends. The tubular shroud is coaxial with the axis of rotation and rotates jointly with the casing when the gas turbine is operated. Between the casing and the tubular shroud a tubus is formed having an annular cross section extending from the gas turbine entry to the exit. In this embodiment the whole annular tubus is rotating when the gas turbine is in operation.

In a first preferred embodiment a number N of compressor blades and an identical number N of turbine blades is provided. In this embodiment each compressor blade merges in a corresponding turbine blade thus forming an individual number of N-1 rotatable passages and accordingly N-1 rotatable combustion chambers each passage and combustion chamber being formed between two neighboring blades. In the section of the N-1 combustion chambers the side walls of these combustion chambers may extent parallel to the axis of rotation.

This design consists of isolated stream tubes, whereby no tip clearance will exist. The numerous blade leading and trailing edges in conventional turbo machines lead to significant losses. This source of loss is minimized in this invention because only one blade leading edge exists at the compressor inlet and one blade trailing edge at the turbine outlet.

In a second preferred embodiment of the invention a number N of compressor blades and a number M of turbine blades is provided. The number N and M may be equal or different. In this first embodiment the combustion chamber may be formed without any blades or side walls between the compressor section and the turbine section. So the thermal loading of the combustion chamber walls can be reduced. In particular hot gas recirculation to the impeller can be avoided by means of thick impeller trailing edges which in effect act as flame holders.

In this construction the impeller separates the entry of the gas turbine into N stream tubes that are merging into one big stream tube in the section of the combustion chamber. The compressor blades are separating the exit area into M stream tubes.

In this embodiment the compressor blades may be converging at their trailing edges asymptotically to a plane being parallel to the axis of rotation. Furthermore in combination with this or any other embodiment a fuel outlet may be disposed in the trailing edge of at least a partial number of all compressor blades, in particular in the trailing edges of all compressor blades. Integrating a fuel outlet in the trailing edge may be simplified if this trailing edge is thicker that the leading edge or an intermediate part of the compressor blade.

Thicker trailing edges are favorable flame holder for the rotating combustion chamber. Thick blade roots may be needed for structural reasons anyway. At impeller exit they are favorably combined with thick trailing edges of the blades. Thick wedge type impeller trailing edges reduce diffusion at the radially directed trailing edges of the impeller and will act as flame holders. The fuel will be injected from thin nozzle systems, which are integrated in the thick trailing edges.

The aforementioned fixation member may be formed by a shaft which is coaxial with the axis of rotation and fixed with respect to an external reference. Such a reference may be the earth or a vehicle. The bearings may be disposed between the shaft and the aforementioned tubular shroud allowing the unitary system to rotate around the shaft. In the presented invention, the fuel supply to the rotating system represents a particular challenge. According to a preferred embodiment of the invention it is possible to charge the fuel from a supply system to the gas turbine with a hollow shaft. When using such a shaft a fuel channel may be provided in the fixed hollow shaft, the fuel being guided by a rotatable sealing system from this fuel channel to fuel nozzles. These nozzles may be positioned in the rotatable tubular shroud, in particular in the section of the combustion chamber/s and/or in the compressor blades, in particular in their trailing edges.

In this embodiment the fuel must be transferred from the stationary hollow shaft via a mentioned sealing system that can be realized through labyrinth seals, simmer rings or brush seals to the supply lines of the rotating combustion chamber.

In another embodiment the tubular shroud may be fixed on a shaft coaxial with the axis of rotation, the shaft being rotatable jointly with the tubular shroud and the casing, in particular the shaft being rotatably supported by bearings on the fixation member. Tubular shroud and the shaft also may be the same element. If the bearings are positioned in a preferred overhung position the bearings may be upstream the casing.

In this embodiment a fuel channel may be provided in the rotatable shaft, the fuel being guided from the fuel channel to fuel nozzles in the rotatable tubular shroud and/or compressor blades as described in the aforementioned embodiment.

If the shaft is rotating and is connected directly with the rotating casing, there is no need for any sealing or bearing system in the hot section of the gas turbine. In this case, the rotating shaft has to be brought through the engine inlet in a stationary fuel reservoir, or the fuel will be brought in the rotating system.

No matter which one of the mentioned fuel supplies is used there is no need for high fuel pressure in the hollow shaft. This is due to the fact that the fuel will be accelerated by the centrifugal force in the rotating fuel lines, whereby a good fuel atomization in the combustion chamber can be guaranteed. The fuel channel may be aligned along the turbine inner shroud thereby pre-heating of the fuel is possible. Furthermore fuel may be alternatively injected through a porous inner shroud of the combustor. All that is driven simply by the centrifugal force.

In another embodiment the fixation member also may be formed by a tubular member coaxially surrounding the rotatable casing and fixed with respect to an external reference, the bearings being disposed between the tubular member and the outer wall of the casing.

Basically the invention's gas turbine concept may be compared to that of a ramjet engine, which operates in a rotating frame of reference. A ramjet has no rotating components, and is nothing more than a pipe with an inlet (diffuser), a combustion chamber and an exhaust nozzle. A ramjet can only be operated if there is already sufficient incoming airspeed at high flight Mach numbers. In effect, the

aerodynamic ram energy at the inlet is converted into pressure. In the combustion chamber the total energy rises through chemical reaction with the added fuel. Finally, the hot gases expand to ambient pressure in the nozzle in order to generate thrust. Basically, a stream tube of the invention may be described as a curved rotating ramjet because every component of the ramjet engine can be found in the invention. The ramjet aerodynamics are generated in the rotating relative system, whereby this gas turbine concept can also be used in stationary applications.

Preferred embodiments of the invention are shown in the figures.

The design of this gas turbine concept is shown schematically in Figure 1.

Compressor blades 1 and turbine blades 2 are fixed with their respective outer radial ends at the inner wall of a casing 3 and with their respective inner radial ends at the outer wall of a tubular shroud 4. Between the compressor blades 1 and the turbine blades 2 and the casing 3 and tubular shroud 3 an annular free space exists forming the combustion chamber 5 in with fuel may be supplied by at least one fuel channel 6 ending in a fuel nozzle 7 positioned in the tubular shroud 4 in the section of the combustion chamber 5. Fuel is supplied to the fuel channel 6 through a hollow shaft 8 that is fixed with respect to an external frame of reference. On the shaft 8 bearings 9 are mounted supporting the tubular shroud 4 und thus the whole unitarily rotatable system of elements 1-7.

Fuel is transported to the nozzle 7 by centrifugal forces when the whole system is rotated. The fuel may be ignited by means of an ignition plug 10 near to the nozzle 7.

In rotation air will enter from the left side and will be compressed by the

compressor blades 1. The compressed air in the combustion chamber will be mixed with injected fuel and combusted. The combusted fuel-air mixture expands and is exhausted through turbine blades 2 at the right exit of the gas turbine thus driving the whole system on the fixed shaft 8.

The combustion chamber 5 may be composed of one single volume or of several stream channels separated by neighboured chamber walls merging with the compressor and turbine blades as described in the general part of this text.

The last compressor stator vanes upstream of the combustion chamber and the first turbine stator vanes downstream of the combustion chamber which are known from common gas turbines are no longer needed in the gas turbine according to this invention due to the fact that the rotating combustion chamber 5 connects the compressor impeller with the turbine rotor.

In Figure 2, a single meridional flow channel of an embodiment having several combustion chambers can be seen. Each compressor blade 1 merges into a respective turbine blade 2 thus forming combustion chamber side walls 5A and 5B surrounding each of the individual combustion chambers 5. The sketched velocity triangles identify the aerodynamic performance of the invention. The compressor section (2)-(3) may be made of a metal, in particular titanium, whereas the combustion chamber 5 or (3)-(4) and the turbine (4)-(5) may favorably be manufactured out of a ceramic material. With titanium compressor blades 2, thin leading edges 1A can be realized, which is an important design feature for high compressor efficiency. Ceramic materials have the advantage of being able to tolerate very high temperatures, which enables a significant increase in the turbine inlet temperature. High turbine inlet temperatures lead to an increase in thermal efficiency and specific thrust. For this reason, in modern high thrust gas turbines, high turbine inlet temperatures are achieved. This is only possible by using extensive cooling technology in the combustion chamber and the high pressure turbine. As mentioned before, the loss intensive and costly cooling technology can be avoided in the turbine and combustion section of this invention in particular when made of ceramic.

In contrast to figure 2, figure 3 shows a meridional flow channel of an embodiment having a single combustion chamber extending between the compressor blades and the turbine blades. There are no side walls 5A or 5B.

As shown in Figure 4 the rotating casing 3 may also be mounted from the outside, whereby the use of larger bearings 9 and improved accessibility would be ensured. With this arrangement, the hollow shaft 8 would only act as a fuel supply and would not have to absorb any external force loads. Because the shaft 8 is connected directly to the rotating core engine, formed by the casing 3 and inner tubular shroud 4 interconnected by the blades, the fuel supply in the rotating system is solved easier than by transferring the fuel from a nonrotating shaft with a sealing system to the rotating combustion chamber 5.

In another (not shown) bearing option a hollow shaft will be connected directly to the rotating core engine and the bearings will be placed in struts, which are located in front of the compressor and downstream of the turbine. This option also allows the use of a rotating hollow shaft with good accessibility in the struts. The fuel supply to the hollow shaft can be made in the IGV's (Inlet Guide Vanes), whereby the fuel shaft will not be routed through the inlet of the aircraft system. Via a rotating connection, the fuel can be infused from a stationary reservoir into the rotating hollow shaft.

As an option, the rotating casing is being held by an overhung bearing mount located upstream of the compressor in the cold region as shown in Figur 4.

Bearings and lubrication system then is solely in the cold region avoiding all (typical) hot temperature problems.

Generally in all embodiments besides the mechanical stress load of the rotating casing, the combustion process in the rotating combustion chamber will be the biggest technological challenge. The fuel droplets will be forced outward towards the casing by the centrifugal force, which could complicate a good fuel atomization and therefore stable combustion. The combustion of hydrogen in the rotating combustion chamber can minimize the problems of liquid droplet combustion, because the density of hydrogen can be regulated over the gas pressure.

Consequently there will be no density difference between the compressed air and the gaseous hydrogen, or any other gaseous fuel, whereby the gaseous fuel can be mixed with air, just like under non buoyant conditions. The characteristic flame velocity of turbulent hydrocarbon flames is about 2 m/s to 6 m/s. The average flow velocities in combustion chambers are all above the mentioned values of the flame propagation velocity. In order to stabilize the flame and ensure continuous combustion, additional geometries in the combustion chamber will be necessary.

In combustion chambers of conventional jet engines, a recirculation zone with low flow velocities and long residence time may be induced in the front part of the combustion chamber. This recirculation zone acts as a result of mass exchange with the surrounding fluid as a source of continuous ignition and combustion of continuously forming fresh fuel-air mixture. The recirculation zone can be generated in different ways. One possibility is to use a flame tube with a swirl generator similar to conventional combustion chambers. With this arrangement a flame might be stabilized, but the flame tube would be highly loaded by

mechanical stresses due to the centrifugal forces. Therefore, a recirculation has to be created without using additional cantilever geometries like a flame tube or swirl generator. The fluid particles at the compressor outlet will be directed to the casing due to the centrifugal force. Thus, a zone of high velocity develops at the casing, whereas a low velocity zone with flow separation and recirculation develops at the hub. This circumstance can be used, by placing the fuel injection at the hub, namely in the low velocity zone. As mentioned before, in conventional combustion chambers the recirculation zone is generated by a swirl generator. Combining a backward facing step with a cavity, as shown in Figure 5, a recirculation zone 11 can be generated because a vortex system will be initialized in the cavity. Such a recirculation zone may be used in any embodiment of the invention.

In this vortex system, in particular the cavity generating this vortex system, the fuel injection may be implemented to guarantee optimal mixing of the fuel-air mixture with the hot reaction products.

As mentioned before it is possible to integrate fuel injection nozzles 7 in the trailing edges 1A of compressor blades 1. Such a flame holder system integrated in thick impeller trailing edges (compared to the leading edges) can be seen in Figure 7.

To reduce the thermal loading of the mechanical highly loaded casing, buoyancy driven casing cooling in combustor, turbine, and nozzle can be integrated. Cold air forms an isolating layer on the outer surface (=casing) of the flow channel. This is due to the large centrifugal forces which drive the heavy cold air outward and the light hot burnt gas inward. Such a temperature distribution can be realized, if the combustion is stabilized at the hub. The cold air stream does not participate in the combustion process and the compression and expansion of this stream therefore effectively represents a loss. However, due to the vaneless and gapless

compression and expansion the associated efficiencies are very high so that the cold air stream loss is small and a comparatively high amount of this cold air stream is acceptable. After the structural design of the invention has been discussed in detail, the possibilities of integrating the gas turbine in overall engine concepts will be presented. The described invention corresponds to a gas turbine engine without bypass. Since such engines are only an adequate solution for very high flight Mach numbers, it will be shown in the following how this gas turbine can be integrated into different overall engine concepts.

Turbofan engines have their optimum operating range at high subsonic flight Mach numbers. An increase of bypass ratio leads at these flight Mach numbers to an improvement in propulsion efficiency η ₃ and thus to increased overall efficiencies η ₉ . In order to offer an optimum engine concept for this flight Mach numbers, the rotating core engine will be integrated in a turbofan. The turbofan concept, shown in Figure 7 is called a Counter Rotating Shrouded Fan, because it has counter rotating fan blades and is covered by a casing.

This engine concept allows either a high bypass ratio at low fan pressure ratios, or high fan pressure ratios at lower bypass ratios. The second option should be favored at higher flight Mach numbers. The core engine with a single-stage centrifugal compressor will be limited to pressure ratio around 5 because higher pressure ratios cannot be realized with single-stage centrifugal compressors. The thermal efficiency can be significantly increased by increasing pressure ratio, wherefore an axial compressor will be placed upstream of the core engine. This booster stage consists of a counter-rotating axial-flow compressor, having the advantage that no stator vanes must be used and compressor work can be done by both compressor blades. Assuming that each axial blade row of the booster stage and the fan stage can realize a pressure ratio of 1 ,6 and the centrifugal compressor can handle a pressure ratio of 5 , the overall pressure ratio of the turbofan engine can be increased to approximately 20.

This pressure ratio already guarantees very competitive thermal efficiencies. In this configuration, the main shaft 8 rotates in the opposite direction to the core engine 1 ,2,3,4,5. On this shaft 8 the power turbine 12, one blade 13 of the axial compressor, and one fan blade 14 is mounted. The axial power turbine 12 is used to propel the first fan blade 14 and the second compressor blade 13. There is no stator vane in the entire turbofan engine, because all stages are counter rotating stages. In Figure 7 the components, which rotate out of the leaf level are

symbolized by a circled dot, the components which rotate in the opposite direction, are symbolized by a circled cross. The bearing system for the rotating shaft can be placed in struts upstream of the fan and downstream of the turbine. The engine nacelle can be integrated under the wing, like conventional turbofan engines.

The distinctive feature between a high bypass ratio turbofan engine and a turboprop engine is that the fan of the turbofan engine is surrounded by a nacelle. At small to medium flight Mach numbers the turboprop engine will achieve the best overall efficiencies. For this reason, it will be shown how the gas turbine according to the invention can be integrated into a turboprop engine.

The configuration outlined in Figure 8 represents only one way of linking the propeller operation with the core engine. In the shown configuration, the propeller 15 acts as a pusher and is attached directly to the power turbine 12. Since the power turbine 12 is working decoupled from the core engine 1-5, this version is similar to a twin shaft concept, although only one shaft is used. The power turbine 12 drives the propeller 15 and one blade row 13 of the counter-rotating axial compressor. The booster compressor stage 13/16 in front of the core engine enables an increase in pressure ratio, which raises the thermal efficiency. To install the propeller 15 directly on the power turbine is a mechanically

sophisticated solution, because high centrifugal forces attack at the turbine blade root.

Another possibility is to mount the propeller 15 on the shaft 8 as a puller in front of the compressor section, which is also the mechanically better solution. In both presented options of the turboprop, the propeller rotates with the rotation speed of the power turbine 12. According to the Eulerian Turbine Equation, the extracted turbine work can be increased either by high circumferential velocities or by strong fluid deflection. Because the maximization of turbine work is the aim of power turbine development, the increase of circumferential velocity by limited

aerodynamic blade loading is the only possibility to achieve this aim.

If the propeller speed is not reduced by a gear, high relative Mach numbers will appear at the propeller tip, which lead to an increase in losses and an increased acoustic loading.

One possibility to relinquish the gearbox is to integrate the core engine in an open rotor concept, whereby two fast counter rotating fan blades without a nacelle will be used. Basically this open rotor concept is nothing else than the counter rotating fan, which was shown in Figure 7, except that the fan blades are not covered by a nacelle.

In a reduction gear as used in conventional turboprop engines, the propeller speed is coupled to the speed of the power turbine. One possibility to avoid this dependence is to run a generator 17 by the power turbine, which supplies an electrical driven fan stage. In this configuration the generator 17 corresponds to an electrical gearbox and is similar to the stationary gas turbine, which is shown in Figure 9.

In this option the gas turbine can be operated at constant revolution speed, as load peaks can be buffered by a battery, which is connected between the generator and the electric motor. Due to this decoupling of the gas turbine speed from the demanded load condition, the loss intensive part load behavior of conventional jet engines can be avoided. In the higher flight Mach number range the turboprop engine will show significant performance degradation at compressor pressure ratios over 12. Best performance is achieved at all flight Mach numbers in the range of compressor pressure ratios around 10. Higher compressor pressure ratios will further reduce the specific fuel consumption, but due to the simultaneous decrease in power output a further increase of compressor pressure ratio is less attractive. With the presented gas turbine with one counter rotating booster stage 16/13, the increase of compressor pressure ratio is limited. Hence, it is beneficial that no high pressure ratios are desired for turboprop engines. Assuming that each axial blade row of the booster stage can achieve a pressure ratio of 1 ,6 and the centrifugal compressor can handle a pressure ratio of 5 the overall pressure ratio of the turboprop engine can be increased to approx..13. As mentioned before, a pressure ratio of this magnitude is sufficient for the optimal operation of a turboprop engines.

A turboshaft engine differs from a turboprop engine by the fact that in the turboshaft engine the complete exhaust energy is consumed in the power turbine and no usable thrust from the exhaust stream is obtained. A turboshaft engine is equivalent to a stationary gas turbine and can be used for electrical power generation. The simplest way to integrate the invention into a stationary gas turbine is to mount a generator 17 directly on the rotating casing 3. This

configuration is shown in Figure 10, which is designed as a single-shaft gas turbine.

Thus, the gas turbine is operated without an upstream booster stage, and no power turbine will be used. For this reason the radial flow turbine has to power the radial compressor and the generator. For radial flow turbines the potential conversion of work is limited to specific enthalpy ratios oi » -1, whereby in this configuration the radial turbine will quickly reach its limit. This problem makes an additional power turbine necessary, as shown in Figure 9. The use of a compressor booster stage 16/13 is advantageous, because the thermal efficiency of stationary gas turbines increases significantly with the compressor pressure ratio.

For a stationary gas turbine of this magnitude a wide range of applications is conceivable. In aviation, the use as a turboshaft engine for helicopters, as an engine for small propeller planes, or as an auxiliary power unit (APU) is conceivable. Auxiliary power units supply the aircraft with electrical and pneumatic power and are used as an autonomous energy source only on ground. For this reason, auxiliary power units just operate temporary. During the flight, the APU represent unused mass for the aircraft system, so a high power density is the primary design feature and low fuel consumption is of secondary importance. The presented invention combines exactly these characteristics, thereby the invention is predestined for the use as APU.

Such a small gas turbine can be used as a stationary gas turbine for external power supply in emergency cases or in conflict areas.

Due to limited battery capacity and limited battery loading speed, electric vehicles nowadays usually have a shorter range than vehicles with internal combustion engines, which makes the use of a range extender required. Today, internal combustion engines or fuel cells are used as a range extender. Because the range extender is only used for long distance drive and will not be operated in urban traffic situations, the high power density of a small gas turbine possibly legitimates the use as a range extender.

The gas turbine with rotating casing can be used as a core engine of a

conventional gas turbine. Either integrated on the same system of shafts or as a stand-alone separate high pressure gas turbine. The highly efficient rotating core (through gapless and vaneless compression and expansion and associated highest aerodynamic efficiencies) is best suited for high-pressure compression which usually suffers from poor efficiencies due to huge radial gaps and large wetted areas. This usually limits the total pressure, especially for small gas turbines where the radial gaps and wetted areas are largest relative to the through flow.

Small gas turbines often have a last stage radial compressor prior to the combustion chamber. This radial stage along with the combustor and high pressure turbine may then be effectively exchanged by the rotating casing configuration of this patent claim.

The proposed rotating casing configuration is even more promising if applied to a standalone high pressure core engine such that the compressed air of a large and standard gas turbine is extracted and fed into a separate one spool rotating casing gas turbine as proposed by this patent claim with small outer diameter. The exit gas of this small rotating casing gas turbine can then be fed back into the low pressure section of the large gas turbine. By doing so, the small rotating casing gas turbine has much larger blading aspect ratios and thereby much less wetted area (and per se no radial gaps) relative to the gas stream through flow and in return much higher efficiencies than the core engine of a large gas turbine which is bound to operate at a large diameter and therefore much more unfavorable conditions. The higher efficiencies of the separate small gas turbine with rotating casing now enables a much higher compression and overall pressure ratios and accordingly much higher thermal efficiencies of the (combined) gas turbine.