DE102019130861B4 - JET ENGINE WITH IMPULSE TURBINE - Google Patents

Abstract

In one aspect, the invention relates to an engine for an aircraft or rail vehicle with a jet engine which drives a pulse turbine. The impulse turbine in turn directly drives a drive unit to generate propulsion for the aircraft or rail vehicle.

Description

In one aspect, the invention relates to an engine for an aircraft or rail vehicle with a jet engine which drives a pulse turbine. The impulse turbine in turn directly drives a drive unit to generate propulsion for the aircraft or rail vehicle.

Background and prior art

Aviation represents the most progressive development in transport. Its growing importance in the global flow of goods is reflected in the rapid development of new aircraft and the rapid expansion or construction of new airports. In the course of aviation, goods could and can also be transported over great distances that otherwise could not be transported (e.g. living organs or perishable food). Rescuing people from dangerous situations has also benefited greatly from aviation.

At present, most aircraft are dependent on large airports and the handling of goods requires a lot of lead time and lead time. Helicopters and VTOL aircraft can be used more flexibly than fixed-wing aircraft, but they consume much more fuel in relation to payload or passengers. Their range is therefore very limited and their range of uses is limited to special operations and a few exceptions in the commercial sector.

From a physical point of view, a vertical take-off aircraft with dynamic lift always consumes more energy than a fixed-wing aircraft with comparable weight and performance parameters that accelerates on a runway. While an accelerating aircraft on the runway converts its propulsion energy almost exclusively into propulsion energy and the lift forces are generated by the acceleration without additional energy requirement, in a vertical take-off or helicopter the propulsion force has to be converted directly into lift force. This counteracts the force of gravity directly, which leads to a much higher energy expenditure.

In the 21st century, a great many airfields and airports are developed in such a way that various types of aircraft can be used, from standard fuselage to wide-body aircraft and various types of cargo aircraft. An extensive network of rail and road routes also enables smaller towns and distribution centers to be opened up by air traffic. Due to the increasing world population and the accompanying steadily increasing flow of goods, it can be assumed that air traffic, especially traffic with fixed-wing aircraft, will increase in the near future. At the same time, however, it is assumed that the demand for crude oil will no longer be adequately met within the next three decades.

In addition to researching new energy sources for aerospace engines, increasing the efficiency of these aerospace engines is a challenge in itself.

Fixed-wing aircraft combine a long range with a high payload and moderate fuel consumption. They currently represent the most economical variant of commercial air transport and, thanks to their high speed, they offer the possibility of transporting perishable or sensitive freight over long distances. For example, transplants of living organs or the import of flowers or food over great distances are possible. In addition, the entire tourism industry benefits from civil aviation, and when it comes to cross-regional travel, motorway, rail and ferry connections in Europe and North America are facing increasing competition from commercial aircraft. At the same time, extensive airport facilities and an extensive airport infrastructure are necessary to maintain ongoing and steadily growing air traffic. Airports also exert a strong influence on the surrounding environment and localities. For runways, miles of aisles must be leveled, the subsurface must be paved and the airspace in front of and behind the runway must not be disturbed up to a certain flight altitude. Since there are not always the spatial possibilities for expanding an airport, direct connections with large-capacity machines are not always possible. This in turn means additional handling time for certain goods or for passengers due to the need for stopovers and reloading operations. Compared to helicopters or airships, aircraft are dependent on external controls. It is used nowadays in rail, sea and air transport and means that the drivers of the respective watercraft, rail or aircraft are dependent on the instructions from control centers located outside their own vehicle in order to ensure safe transport operations . Railway trains own at A long braking distance at high speeds, so that when the brakes are actuated, the train only comes to a stop outside of the area that the driver can see. Aircraft and ocean-going vessels have a high degree of inertia in inputting controls and therefore there is a need for them to communicate the position of other aircraft or ships in order to avoid collisions. These processes are supported by radar systems and satellite navigation or, in rail traffic, "control" is carried out by signal systems. In contrast to this, one speaks of internal control in motor vehicle, bus and truck traffic. The control of the vehicle is the responsibility of the driver. The traffic is influenced from outside by means of traffic control systems and traffic rules, the route and speed are selected by the driver. In air traffic, a distinction is made between IFR (instrumental flight rules) and VFR (visual flight rules) flights, which is roughly equivalent to the distinction between internal and external control. In the field of private and hobby aviation as well as with helicopters, there is the possibility of applying VFR flight rules in everyday flight operations due to the high maneuverability and the lower speed compared to jet-powered commercial aircraft. In commercial aircraft, where this option is also used under certain circumstances, it only plays a very minor role at the present time due to the high traffic density.

With the development of heavy-duty helicopters such as the Mil Mi-26, the transport capacities of small and medium-sized cargo aircraft such as the C-160D can be covered, whereby the need for specially created airfields is no longer necessary or flight operations between different airfields can be ensured with a helicopter a fixed-wing aircraft of comparable performance would not be suitable.

In order to counter the problem of poorly developed airfields or unpaved runways, various aircraft were tested that only require short take-off and landing routes and also have low minimum speeds. This ability is known as the STOL ability ("Short Take Off and Landing" ability). Prominent examples are military transport aircraft such as the Transall C-160D or the American C-130 Hercules. The development of aircraft types that can be used outside of airfields began historically before the first helicopters were ready for series production. It is therefore understandable that their field of application partly covers that of modern helicopters. Since the advancing development of helicopters means an increasingly important alternative to transport aircraft, modern developments such as the A400M from Airbus or the Chinese Shaanxi Y-9 show the attempt to increase the speed, the higher flight altitude and the lower fuel consumption of fixed-wing aircraft with the flexibility of helicopters to combine. Small transport aircraft such as the Polish An-28 require a take-off distance of around 270 m. A C-160D requires an average take-off distance of 660 m and can also operate on unpaved runways. 250 - 700 m describe a relatively short runway length (a large aircraft such as an A330 requires about 2000 m runway length) and it can be implemented on large, clear cross-country roads. During the Cold War, various motorway sections in the FRG were upgraded for temporary air traffic with military aircraft. The ability to take off and land on unpaved runways also extends the range of applications for transport aircraft. Another field of application is served by amphibious aircraft. Nowadays they are used in fire fighting and occasionally also in commercial traffic in the Pacific or the Caribbean. In addition to individual helicopter models that can also land in water, they are able to connect two islands if landing on land is not possible. The Russian manufacturer Beriev, for example, advertises its Be-200ChS model for precisely this purpose.

Another aspect is that, according to theoretical considerations, fixed-wing aircraft can be better designed to transport very large payloads. Aircraft like an A380-800 from Airbus or the Ukrainian Antonov An-124 can carry up to 80 tons of payload or more. Constructing helicopters that can transport more than 40 tons of payload is very difficult if not impossible. While the wings of aircraft can have a high lift surface and large wingspan, this is hardly possible with rotor blades of helicopters. In the largest and heaviest helicopter ever built, the Soviet Mi-12, two main rotors were necessary to generate the necessary lift forces. With the rotors running, it had a span of 67 m and a payload of 40 tons. In contrast, cargo aircraft with a comparable span usually have a payload of up to 70 tons.

In theory, helicopters can currently already relieve the heavily subsidized truck traffic in shuttle operation between large logistics centers. Their range and loading capacity are comparable to the average semi-trailer truck. This idea is already being tested to a small extent in the form of the cargo drone. Only the high fuel consumption of helicopters is a hindrance to implementation this concept. Replacing helicopters in this application example with small transport aircraft includes the connection to a nearby airfield or the establishment of a short runway.

The two aircraft types PZL An-28 and Mil Mi-171 offer a very good comparison between the concept of the helicopter and that of the fixed-wing aircraft. These two aircraft have almost the same range, payload, cruising speed, maximum altitude, wingspan and cargo hold dimensions. They only differ in fuel consumption, which is much higher with the Mi-171, the total weight, the empty weight and the required runway length, which is practically non-existent with the Mi-171. On the basis of these features it can be concluded that from an economic point of view the advantage lies clearly with the aircraft concept, while from a logistical point of view the helicopter has a clear advantage.

Taking into account the evolving air traffic infrastructure, the steadily increasing global flow of goods and recurring social change, it is therefore important to sound out the flexibility of air traffic. Taking into account various aspects with regard to economy, flight performance and cost economy, the development of the various aircraft types of helicopters and fixed-wing aircraft will result in an increasing convergence of both aircraft variants with one another. This means that in the future helicopters will probably have similar flight performance and a similar economic efficiency as fixed-wing aircraft, while fixed-wing aircraft will adapt more and more to the helicopter concept in terms of flexibility and short take-off capability. However, physical conditions prevent a complete adaptation, so that both aircraft types will still be used in the future. It is therefore important to consider both aircraft types in terms of their economy and product improvement.

The pamphlet US 2018/0051627 A1 discloses a gas turbine engine for an aircraft. In this regard, the aircraft has a first engine part with a compressor and a combustion chamber in which air is mixed with fuel and ignited in order to generate an energy-rich gas flow. This gas flow expands in a turbine section comprised of the first engine part and is then passed to a further exposed turbine via a duct. The exposed turbine is configured to be driven by the gas flow from the first engine part and to generate the thrust for the aircraft.

Out EP 2 966 266 A1 and WO 2015/134081 A2 two-part engine systems are known. A first jet engine is connected to a downstream, exposed turbine via a duct, so that the mass flow from the jet engine drives the turbine.

The document US 10 287 991 B2 discloses a gas turbine engine having a gas generator rotating along a first axis of rotation with at least one compressor rotor, a gas generator turbine rotor and a combustion section. A fan drive turbine rotates along a second axis of rotation, positioned downstream of the gas generator turbine rotor. The fan drive turbine drives a pair of shaft members.

The revelation US 10 710 738 B2 describes an auxiliary power unit with an intercooler for an aircraft. In a preferred variant, the turbine section of the auxiliary power unit comprises a pulse turbine.

Object of the invention

It is an object of the invention to provide an engine for aircraft and rail vehicles without the disadvantages of the prior art. In particular, it is an object of the invention to provide an engine which functions particularly efficiently and conserves resources and which has good to improved performance. At the same time, the design of the engine should be particularly flexible, with individual components being able to be arranged differently from one another without major losses occurring in between during the power or power transmission.

Summary of the invention

• - mindestens ein erstes Strahltriebwerk
• - mindestens eine Impulsturbine umfassend mehrere auf einer Welle gelagerten Schaufeln,
• - mindestens eine Antriebseinheit zur Erzeugung eines Vortriebs des Luft- oder Schienenfahrzeugs,

wobei das erste Strahltriebwerk einen Massenstrom eines Fluides erzeugt, welcher mit einer Strömungsrichtung innerhalb einer Rotationsebene der Impulsturbine auf deren Schaufeln gerichtet ist, wobei die Welle der Impulsturbine mit der Antriebseinheit verbunden vorliegt für einen Antrieb der Antriebseinheit durch die Impulsturbine.In one aspect, the invention relates to an engine for an aircraft or rail vehicle, comprising:

• - at least one first jet engine
• - At least one impulse turbine comprising several blades mounted on a shaft,
• - at least one drive unit for generating propulsion of the aircraft or rail vehicle,

wherein the first jet engine generates a mass flow of a fluid which is directed with a flow direction within a plane of rotation of the impulse turbine on the blades thereof, the shaft of the impulse turbine being connected to the drive unit for driving the drive unit by the impulse turbine.

A jet engine preferably comprises a gas turbine. In particular, a jet engine generates a recoil effect on the basis of a generated air and / or exhaust gas flow. The jet engine preferably comprises a compressor, a combustion chamber and / or a turbine and preferably a thrust nozzle. Air that is sucked in is preferably compressed in the compressor, the fuel is preferably injected and ignited in the combustion chamber, whereby the mixture is preferably accelerated. The turbine, which is preferably connected downstream, is preferably driven by the accelerated mixture and in turn drives the compressor and the fan, which is preferably enclosed. The fan preferably generates the so-called sheath flow. Thus, preferably an air flow is referred to, which is guided to the rear between the actual gas turbine and the outer engine cowling and blown out. The gas preferably expands in the thrust nozzle arranged at the rear end of the jet engine, the flow velocity preferably being increased.

A pulse turbine is preferably a turbine in which the working medium has the same static pressure before and after the impeller. In this case, usable work is preferably generated essentially exclusively from the conversion of the kinetic energy of the working medium. This results in particular from the dynamic pressure of the working medium according to Bernoulli's energy equation. A positive pressure turbine preferably has a degree of reaction of essentially a maximum of 0.5 and in particular essentially zero. The degree of reaction preferably designates a dimensionless number which can assume a value between zero and one and describes the ratio of the enthalpy converted at the rotor to the total enthalpy converted in the turbine. “Rotor” preferably describes the sum of all impellers of a turbine, which are connected to a shaft of the turbine. The impeller preferably designates a rotating part of the turbine, which extracts work from the flowing working medium and preferably comprises a plurality of blades. “Tail unit” designates, preferably, but not necessarily encompassed by the turbine, stationary, vane-like elements which preferably cause the flowing medium to swirl. A tail unit and the associated impeller together preferably form a “stage”. It can be preferred that the turbine comprises at least one tail unit. It can also be preferred that the turbine does not include a tail unit. It can be preferred that the turbine comprises at least one impeller. In the case of a multi-stage turbine, several running wheels (and possibly tail units alternately) are preferably connected in series. The turbine is preferably single-stage. However, the turbine can also comprise several stages.

In the case of a pulse turbine, the degree of reaction is preferably essentially 0, which means in particular that no conversion of enthalpy takes place in the impeller. Only in a possibly enclosed tail unit can essentially the entire static pressure be converted into kinetic energy. In the case of a pulse turbine, essentially the entire enthalpy gradient is converted into flow energy in the possibly (if present) at least one tail unit, the static pressure in the impeller preferably remains essentially constant.

Terms such as essentially, approximately, etc. preferably describe a tolerance range of less than ± 20%, preferably less than ± 10%, particularly preferably less than ± 5% and in particular less than ± 1%. “Similar” preferably describes sizes that are approximately the same. Partly or “approximately” describes preferably at least 5%, particularly preferably at least 10%, and in particular at least 20%, in some cases at least 40%.

Examples of impulse turbines are Pelton turbines, Curtis turbines, Laval turbines and / or through-flow turbines.

With the same power output, a constant pressure turbine rotates preferably 0.707 times (1 / √2) slower than a so-called overpressure turbine with a degree of reaction of 0.5, provided that the mass flow and turbine diameter are the same. At the same speed, a constant pressure stage advantageously emits twice as much power as an overpressure turbine with one for overpressure turbines exemplary degree of reaction of 0.5 provided that the mass flow rate and the turbine diameter are the same.

An impulse turbine preferably comprises a plurality of blades mounted on a shaft. A blade is preferably an individual component of an impeller of the turbine, which is preferably configured to generate a rotation of the shaft, which is preferably connected to the blade, from the kinetic energy of the working medium. For this purpose, the blade is shaped in particular to absorb the kinetic energy and / or the transmitted impulse of the moving working medium.

A drive unit for generating propulsion of the aircraft or rail vehicle can, for. B. be a wheel drive in a rail vehicle or a propeller or jet engine in an aircraft or a rotor in a helicopter.

In particular, the first jet engine generates a mass flow of a fluid, e.g. B. an air / exhaust gas flow. The jet engine and / or a nozzle transporting the mass flow is preferably arranged tangentially to the circumference of the impeller and / or to a circular path of a rotating point of the impeller. This is preferably directed with a flow direction within a plane of rotation of the impulse turbine on its blades. The flow direction is in particular the essential and / or averaged flow direction of the mass flow, which is essentially directed and has an essentially straight alignment without deflection elements. The plane of rotation is in particular the plane in which the blades of the impulse turbine rotate. The shaft to which the blades are preferably attached and with which they rotate is preferably a normal to this plane of rotation. The flow is preferably directed towards the blades in at least one point or region of the impulse turbine in such a way that they are driven with the highest possible efficiency.

In particular, the shaft of the impulse turbine, which is connected to the drive unit, is used to drive the drive unit by the impulse turbine. This connection can include both a direct transmission of the driving force, e.g. B. via a common shaft of impulse turbine and drive unit, as well as indirectly via an interposed gear. A drive of the drive unit by the impulse turbine refers in particular to a mechanical power transmission from the impulse turbine to the drive unit via force-transmitting elements, e.g. B. shafts and / or gears.

In this way, a drive can be provided which has a surprisingly high efficiency for those skilled in the art. Furthermore, the drive offers the possibility of providing the power generated by a jet engine very flexibly and advantageously without the interposition of a gearbox to the drive unit. In this case, deflection elements of the mass flow can preferably also be used, so that the jet engine, impulse turbine and / or drive unit do not have to be arranged in relation to one another in a certain way.

In a preferred embodiment of the invention, the blades of the impulse turbine are configured for a transmission of a mass flow impulse of the fluid, which has a direction essentially within the plane of rotation of the impulse turbine. In particular, the blades have a shape that is favorable in terms of fluid dynamics, in order to ensure the highest possible impulse transmission and to make the engine efficient.

In a further preferred embodiment of the invention, the blades have an essentially perpendicular surface to the plane of rotation of the impulse turbine. This has turned out to be a particularly preferred embodiment of the impulse turbine.

In a further preferred embodiment of the invention, the impulse turbine comprises a turbine of the Pelton type. In the Pelton turbine, the fluid preferably flows in a mass flow (preferably a jet) at high speed from at least one nozzle arranged essentially tangentially to the circumference of the impeller on its blades. Each blade is preferably divided into two approximately hemispherical half-blades by a sharp edge. In the middle of the edge, the mass flow of the fluid hits preferably tangentially. The half-blades in particular have the function of deflecting the fluid essentially in the opposite direction so that the kinetic energy is transferred to the impeller according to the principle of action and reaction.

The peripheral speed of the blades should preferably correspond essentially (half) to the speed of the fluid flow. The fluid flow preferably transfers almost all of its energy to the blades. This is particularly efficient.

In a further preferred embodiment of the invention, the first jet engine is arranged in a plane parallel to the plane of rotation of the impulse turbine or within the plane of rotation of the impulse turbine. The mass flow can then be directed onto the blades, in particular, essentially without losses due to deflection elements that are otherwise required.

In a further preferred embodiment of the invention, the shaft of the impulse turbine is connected directly to a shaft of the drive unit. In this way, an essentially loss-free drive can be implemented.

In a further preferred embodiment of the invention, the shaft of the impulse turbine is connected to a shaft of the drive unit via a gear, in particular a spur gear and / or multi-stage gear.

A spur gear is characterized in particular by a plurality of spur gears on parallel axes, with the axes being able to be drive shafts. In a simple design, a single-stage spur gear has two shafts on each of which a gear is seated, the gears preferably driving each other and determining the translation.

In the case of a multi-stage transmission, the transmission ratio is preferably variable in stages.

Thus, the drive unit can preferably be operated independently of the speed of the impulse turbine. Furthermore, the rotation of the shaft of the impulse turbine can be deflected particularly easily and also transmitted via angles.

In a further preferred embodiment of the invention, the impulse turbine is arranged in a turbine chamber which has at least one inlet area for the mass flow of the first jet engine and at least one exhaust gas opening and which is preferably configured for efficient guidance of the mass flow onto the blades of the impulse turbine. The turbine chamber and / or the inlet area and / or the exhaust gas opening are preferably designed according to fluid-mechanical aspects and allow the mass flow to be guided accordingly. In particular, the turbine chamber has the effect that the mass flow is directed along and / or onto the blades of the turbine and does not flow past the impulse turbine.

In a further preferred embodiment of the invention, the impulse turbine drives a rotor and / or a propeller as a drive unit, the plane of rotation of the impulse turbine being arranged parallel to a plane of rotation of the rotor and / or the propeller and the impulse turbine and rotor and / or propeller having a common Have axis of rotation. The power can be transmitted between the impulse turbine and the rotor / propeller, preferably directly via a common shaft and / or via an interposed gear. In this way, a propeller and, in particular, a rotor of a helicopter can be driven particularly easily and efficiently.

In a further preferred embodiment of the invention, the first jet engine is arranged in a plane parallel to the plane of rotation of the impulse turbine or within the plane of rotation of the impulse turbine. The overall shape of the jet engine is approximately cylindrical and / or essentially rotationally symmetrical, with what is meant here in particular that the longitudinal axis of the jet engine, the cylinder axis and / or the axis of rotation are in a plane parallel to the plane of rotation of the impulse turbine and / or within the Plane of rotation is arranged. The jet engine and / or a nozzle transporting the mass flow is preferably arranged tangentially to the circumference of the impeller and / or to a circular path of a rotating point of the impeller.

With this structure, the mass flow of the jet engine can be directed particularly easily and efficiently onto the blades of the impulse turbine.

In a further preferred embodiment of the invention, the impulse turbine drives a propeller as a drive unit, the plane of rotation of the impulse turbine being arranged parallel to a plane of rotation of the propeller, the impulse turbine and propeller having a common axis of rotation, the first jet engine is preferably arranged in a plane perpendicular to the plane of rotation of the impulse turbine in the direction of flight and the turbine chamber is configured to deflect the mass flow so that it is directed with a flow direction within the plane of rotation of the impulse turbine on its blades. The power can be transmitted between the impulse turbine and the propeller, preferably directly via a common shaft and / or via an interposed gear. In particular, the axes of rotation of the propeller and impulse turbine and the longitudinal axis of the jet engine are arranged essentially parallel to one another and to the longitudinal axis of the aircraft, so that the jet engine can be arranged in the aircraft fuselage in an aerodynamically and fluid-mechanically favorable manner (the inlet of the engine is preferably arranged in the direction of flight). In this case, however, the mass flow at the outlet of the jet engine is also preferably directed in the longitudinal direction, which is why the turbine chamber should enable the mass flow to be deflected in order to direct it with a flow direction within the plane of rotation of the impulse turbine on its blades. Such a deflection takes place preferably with the help of deflection elements, in particular by means of arched tubes for directing the mass flow.

This is a particularly preferred embodiment for an aircraft engine due to the spatial conditions or circumstances in which the engine is accommodated. In contrast to an engine of a helicopter, an arrangement is often not desired for aerodynamic reasons in which the longitudinal axis of the jet engine is not arranged essentially parallel to the longitudinal axis of the aircraft.

In a further preferred embodiment of the invention, the engine comprises a jacketed engine fan, the impulse turbine driving the engine fan as a drive unit, the plane of rotation of the impulse turbine being arranged parallel to a plane of rotation of the engine fan, the impulse turbine and engine fan having a common axis of rotation, the first jet engine and / or a nozzle transporting the mass flow is preferably arranged in a plane perpendicular to the plane of rotation of the impulse turbine in the direction of flight and the turbine chamber is configured to deflect the mass flow so that it is directed towards its blades with a flow direction within the plane of rotation of the impulse turbine.

The power can be transmitted between the impulse turbine and the fan, preferably directly via a common shaft and / or via an interposed gear. As in the example mentioned above, the most aerodynamic and fluid dynamically favorable arrangement of all engine components is in the foreground. The arrangement and / or deflection is preferably analogous to the aforementioned embodiment. The engine fan is preferably arranged with its longitudinal axis in the longitudinal direction of the aircraft.

In the engine housing / the casing or at a suitable point of the aircraft there is preferably a first thrust nozzle / a first jet engine that generates a mixture from inflowing air and supplied fuel, accelerates it strongly, which then drives the impulse turbine, which then via a gear drives the fan, which then provides the actual propulsive power for the entire engine construction.

In contrast to conventional jet engines, there is preferably no core flow generated in this one - instead of the combustion chamber in conventional jet engines, the impulse turbine is preferably located here (although the exhaust gas flow from the impulse turbine can preferably also be ejected from the engine in the reverse direction from the engine, since this is the basic mode of operation Engine does not provide a core flow, the exhaust gas flow from the impulse turbine was preferably not considered as a core flow).

If the (first) auxiliary jet engine provided for generating the fuel mixture is located inside the engine, the inflowing air mass first enters the engine through the fan. Part of this incoming air mass is fed into the integrated first jet engine or the thrust nozzle and is preferably used to generate the mixture, while the far greater part of the air mass is preferably compressed and accelerated by the fan and finally ejected like a recoil to generate the propulsive force becomes.

In a further preferred embodiment of the invention, the turbine chamber and / or engine chamber comprises an arcuate deflection of the mass flow. In this way, the first jet engine and the impulse turbine can in particular be arranged individually with respect to one another, in that the mass flow is deflected as far as possible without loss. If this text speaks of a jet engine without an addition, the first jet engine is preferably meant. The (first) jet engine can preferably be arranged in this way be that it is arranged aerodynamically and / or fluid-dynamically particularly favorable. The deflection is preferably also referred to as a deflection element.

In a further preferred embodiment, the turbine chamber is arranged within the casing of the engine fan and is configured for a hermetic separation of the mass flow from a fluid flow within the engine fan. A large part of the engine can thus preferably be arranged in a space-saving and efficient manner within the casing of the engine fan, wherein the fluid flows of the first jet engine and engine fan can be separated from one another. The first jet engine can preferably be arranged inside or outside the turbine chamber and the mass flow of the first jet engine can be fed to the turbine chamber through deflection elements.

In a further preferred embodiment of the invention, an intake port for air of the first jet engine is arranged on a hub of the engine fan. The first jet engine can thus be integrated particularly well into the aircraft and / or the casing of the engine fan and does not require any additional intake ports elsewhere.

The construction of IRS engines offers, when optimally designed, a high potential for savings in the fuel consumption of various aircraft. On the one hand, this leads to a greater range of the aircraft and, on the other hand, to a higher payload with the same range, since less fuel has to be refueled with the same flight performance.

For helicopters, there are therefore new, increasingly important possible uses. In the event of accidents on seas and oceans, far away from the nearest coast, or in vast desert areas, the time factor in saving human lives plays a special role. Rescue ships (are only) need a lot of time to reach the area of operation, fixed-wing aircraft reach the area of operation very quickly, but in most cases cannot land and rescuing people from distress at sea is therefore a major logistical challenge. Helicopters with a range of Such areas of application can reach 5000 - 6000 km in a few hours and rescue people faster than airplanes or ships.

Another application scenario is, for example, the supply of people in remote areas or disaster areas: In devastated earthquake areas, the supply of the population, which is often cut off from the outside world, is often dependent on clearing roads or repairing the local airport. The supply by helicopter is also complicated nowadays, as these are limited in their range and payload. If helicopters were equipped with an IHT engine, it would be possible for them to carry larger payloads over a greater range deeper into the disaster area.

As a link between helicopters and jet aircraft, turboprop machines are now the preferred means of transport in the commercial medium-haul sector. Due to their higher savings potential than with standard fuselage jet aircraft, they can offer a good alternative to the same when using a TPI engine. In the medium-haul range (is closed), the increased flight time is not very important due to the lower cruising speed.

In the commercial long-haul sector, jet-powered wide-body aircraft dominate at present and very likely also in the future. In the long-haul range and also in the ultra-long-haul range, the compromise between payload and required fuel plays a special role. For this reason, stopovers are made on many flight routes, such as the Tahiti-Paris rotation, in order to reduce the fuel required. Using the example of an A330, it can be shown that the use of an STI engine can reduce the fuel consumption of long-haul aircraft by up to 60% (see p. 37). This allows most long-haul aircraft to fully utilize their payload capacity without sacrificing their range or cruising speed. Depending on the aircraft, this also reduces the actual take-off weight and therefore allows an additional reduction in fuel consumption or use at airports with shorter take-off and landing runways.

Consideration of this problem will be of increasing importance for the future of air traffic. The increasing scarcity of raw materials, in particular the falling crude oil deposits, as well as the steadily increasing pollution of the environment by high exhaust gas emissions require efficient, short to medium term implementable measures to reduce fuel consumption while maintaining the same current transport and traffic volume. IRS engines provide a basis for the development of new, efficient aircraft engines.

Detailed description of the invention and examples

"IRS" engine is the name for an engine in which the core is a combination of a special jet engine or a nozzle and an impulse turbine (IRS = impulse turbine rotary jet engine). A pulse turbine, also called a constant pressure turbine, is a turbine in which the fluid has the same static pressure before entering the turbine and after exiting the turbine. The term "IRS engine" refers to a class of different engines that are divided into subclasses with regard to their intended use. The subclasses differ in helicopter engines (IHT engines), propeller and jet engines (TPI and STI engines).

An IRS engine consists of two different engine systems - an upstream nozzle or a special jet engine that generates a mass flow, and a subsequent impulse turbine that is set in rotation by the mass flow generated by the nozzle and then operates the actual drive unit ( e.g. the main rotor of a helicopter, the propellers of a turboprop aircraft, the engine blades of a jet aircraft or the wheels of a locomotive).

In principle, a pulse turbine works like a paddle wheel in a water mill. The mass flow hits the blades of the turbine and sets them in rotation. The direction of flow of this mass flow and the plane of rotation are flat in one plane. In contrast to a free-wheeling turbine in a conventional aircraft engine, the mass flow hits the blades at the edge of the turbine. It has a larger effective rotation diameter than a free-wheeling turbine in a conventional aircraft engine and is therefore able to transmit a greater torque. In addition, with the help of smaller jet engines or nozzles, a higher flow velocity of the mass flow can be generated than in a conventional engine. In conventional engines, a mass flow is generated in the turbine, which is used to drive a free-wheeling turbine, which drives the propulsion unit of the aircraft via an adjoining gearbox. For example, the main rotor of a helicopter is driven via a bevel gear, the propeller of an aircraft or the engine blades of a jet aircraft are driven via a spur gear. The mass flow is directed through the turbine radially to the shaft. At this point there is a decisive disadvantage of positive pressure turbines compared to pulse turbines: compared to positive pressure turbines with the same dimensions, pulse turbines have a higher power output at the same speed.

• Wird der Massenstrom radial durch die Turbine geleitet, so trifft dieser senkrecht auf die Rotationsfläche der Turbine auf. Bei konventionellem Luftfahrttriebwerken wird, abgesehen von einer kleinen Fläche um die Nabe, die gesamte Rotationsfläche der Turbine durchströmt, was bedeutet, dass der Massenstrom bei jeder einzelnen Triebwerksschaufel eine linienförmige Belastung über ihre gesamte Länge aufträgt.
• Die Linienlast auf jeder einzelnen Triebwerksschaufel lässt sich durch eine punktförmige resultierende Kraft ersetzen, die auf halber Länge bzw. im Bereich von 50% bis 70% des Radius an jeder Triebwerksschaufel angreift. So lässt sich der Massenstrom als Summe von Punktlasten, die zu gleichen Teilen an jeder Triebwerksschaufel im Bereich von 50% bis 70% ihres Radius wirken, beschreiben. Eine genaue Berechnung des Angriffspunktes der resultierenden Kraft ist möglich, er variiert jedoch zwischen verschiedenen Triebwerkskonstruktion, und eine exakte Beschreibung der Lage des Punktes auf der Triebwerksschaufel ist an dieser Stelle nicht von Bedeutung. Entscheidend für die weitere Betrachtung ist die Tatsache, dass der Angriffspunkt der resultierenden Kraft sich nicht am äußeren Rand der Triebwerksschaufel befindet und somit nicht ihr voller Radius ausgenutzt werden kann. Daraus ergibt sich, dass die Resultierende Kraft des Massenstroms an einem geringeren tatsächlich wirksamen Radius angreift und daher auch ein geringeres Drehmoment in der Nabe erzeugt, als wenn die Resultierende Kraft am Rand der Rotationsfläche angreifen würde. Ein weiterer Aspekt, der bei radialer Massenstromeinleitung zu geringerer Leistungsabgabe der Turbine führt, besteht darin, dass die Rotationsfläche senkrecht zur Strömungsrichtung steht. Die Flächen der Triebwerksschaufeln sind dabei angestellt zur Strömungsrichtung, so dass bei Auftreffen des Massenstroms Axiale Kräfte entstehen, die die Turbine in Rotation versetzen.
• Dieses Phänomen wird als Reaktionsgrad beschrieben. Da die Wirkungsrichtungen von radialen und axialen Kräften senkrecht zu einander stehen, wird bei Turbinen dieser Bauart meist ein Reaktionsgrad von 0,5 realisiert.

This is explained as follows:

• If the mass flow is passed radially through the turbine, it hits the rotating surface of the turbine perpendicularly. In conventional aircraft engines, apart from a small area around the hub, the entire rotating surface of the turbine flows through, which means that the mass flow applies a linear load to each individual engine blade over its entire length.
• The line load on each individual engine blade can be replaced by a point-like resultant force that acts on each engine blade at half the length or in the range of 50% to 70% of the radius. The mass flow can be described as the sum of point loads that act equally on each engine blade in the range of 50% to 70% of its radius. An exact calculation of the point of application of the resulting force is possible, but it varies between different engine designs, and an exact description of the position of the point on the engine blade is not important at this point. The decisive factor for further consideration is the fact that the point of application of the resulting force is not located on the outer edge of the engine blade and therefore its full radius cannot be used. This means that the resultant force of the mass flow acts on a smaller actually effective radius and therefore also generates a lower torque in the hub than if the resultant force were to act on the edge of the surface of rotation. Another aspect that leads to a lower power output of the turbine when the mass flow is introduced radially is that the surface of rotation is perpendicular to the direction of flow. The surfaces of the engine blades are positioned in relation to the direction of flow, so that when the mass flow hits, axial forces arise that set the turbine in rotation.
• This phenomenon is described as the degree of reaction. Since the directions of action of radial and axial forces are perpendicular to each other, turbines of this type usually achieve a degree of reaction of 0.5.

These phenomena do not occur in this form in impulse turbines. The introduction of the mass flow takes place perpendicular to the axis of rotation, ie the direction of flow lies flat in the plane of rotation. As a result, the rotation-generating force component of the mass flow develops a significantly greater effect than in the case of an overpressure turbine with a degree of reaction of 0.5. In addition, the mass flow attacks the edge of the surface of rotation. The torque thus generated in the hub is generated over the entire length of the radius and thus reaches a value that is up to twice as high as if the mass flow would only act on half the radius. This explains why constant pressure turbines at the same speed have a power output that is up to twice as high as overpressure turbines with the same dimensions. Conversely, this means that compared to an overpressure turbine, less power is required to drive a constant pressure turbine. A pulse turbine, as used in an IRS engine, has a similar shape to a Pelton turbine. A Pelton turbine is a modern water wheel design. Comparable to a known water mill wheel, in which the introduced water is not additionally accelerated, a Pelton turbine is driven by a water flow that moves in the plane of rotation. The principle can be transferred to the IRS engine. _{An efficiency η turbine} of 90%, comparable to a Pelton turbine, is assumed for the impulse turbine.

(Quelle: https://www.easa.europa.eu/sites/default/files/dfu/EASA%20E110%20TCDS%201ssue%207%20LEAP-1A-1C.pdf; 23.10.2018, 21:45 Uhr) Die maximale Drehzahl ergibt sich aus dem höchstmöglichen Drehmoment, der direkt in der Impulsturbine wirken kann. Dieser resultiert aus den Fliehkräften, die vom verwendeten Werkstoff und dessen Gewicht und Dichte abhängig sind, und der Drehzahl. Die verwendeten Werkstoffe sind für die jeweilige Impulsturbine nicht variabel, wohl aber die Drehzahl. Auf Grund der Impulsturbine in ihrer Eigenschaft als Luftfahrt-Antriebs-Komponente wird davon ausgegangen, dass die verwendeten Materialen Drehzahlen in Höhe von $19391\frac{U}{min}$

zulassen. Die maximale min Drehzahl besitzt maßgeblichen Einfluss auf die Strömungsgeschwindigkeit in der Impulsturbine. Beide Faktoren stehen in wechselwirkungsvollem Zusammenhang und zugleich wird für beide Faktoren ein höchstmöglicher Wert angestrebt. Ähnlich wie bei der Abschätzung der maximal möglichen Drehzahl lässt sich der maximale Wert der Strömungsgeschwindigkeit abschätzen. Aus der Überlegung, dass die erzeugte Schubkraft das Produkt aus Massenstrom ṁ_T und der Strömungsgeschwindigkeit v_s darstellt, lässt sich ableiten, dass mit zunehmender Strömungsgeschwindigkeit v_s der benötigte Massenstrom abnimmt und sich damit auch der Treibstoffverbrauch reduziert. Dabei spielen folgende Überlegungen eine Rolle: Zum Einen wird die Festigkeit des verwendeten Werkstoffs einbezogen, zum Anderen die Strömungsgeschwindigkeit in anderen Systemen. Bei Raketentriebwerken, wie sie in der Raumfahrt verwendet werden, lassen sich Austrittsgeschwindigkeiten von bis zu $\mathrm{3000,0}\frac{m}{s}$

realisieren. Der austretende Massenstrom erzeugt jedoch lediglich einen Impuls im Triebwerk und dient nicht dazu, einen weiteren Körper durch Auftreffen zu beschleunigen. Die höchstmögliche Strömungsgeschwindigkeit v_s wird durch das verwendete Material in der Impulsturbine bedingt. Ist die Geschwindigkeit zu hoch, ist davon auszugehen, dass die Impulsturbine beschädigt wird. Von entscheidender Bedeutung ist daher die Festigkeit des verwendeten Materials in der Impulsturbine.A determining factor in an IRS engine is the speed of the impulse turbine. In order to be able to estimate a maximum value here, it makes sense to orientate yourself on the aviation engines currently in use. In a modern CFM LEAP engine, the maximum speed is included in the high-pressure compressor $19391\frac{U}{min}.$

(Source: https://www.easa.europa.eu/sites/default/files/dfu/EASA%20E110%20TCDS%201ssue%207%20LEAP-1A-1C.pdf; 23.10.2018, 9:45 p.m.) The maximum speed results from the highest possible torque that can act directly in the impulse turbine. This results from the centrifugal forces, which depend on the material used and its weight and density, and the speed. The materials used are not variable for the respective impulse turbine, but the speed is. Due to the impulse turbine in its capacity as an aviation drive component, it is assumed that the materials used have speeds of $19391\frac{U}{min}$

allow. The maximum rpm has a significant influence on the flow speed in the impulse turbine. Both factors are interrelated and at the same time the highest possible value is sought for both factors. Similar to the estimation of the maximum possible speed, the maximum value of the flow velocity can be estimated. From the consideration that the generated thrust represents the product of the mass flow ṁ _T and the flow velocity v _s , it can be deduced that _{the required mass flow decreases with increasing flow velocity v s} and thus the fuel consumption is also reduced. The following considerations play a role: On the one hand, the strength of the material used is included, and on the other hand, the flow velocity in other systems. With rocket engines, such as those used in space travel, exit speeds of up to $3000.0\frac{m}{s}$

realize. However, the emerging mass flow only generates an impulse in the engine and does not serve to accelerate another body by hitting it. The highest possible flow velocity v _s is determined by the material used in the impulse turbine. If the speed is too high, it can be assumed that the impulse turbine will be damaged. The strength of the material used in the impulse turbine is therefore of crucial importance.

Die Auftrittsfläche des Massenstroms entspricht der Fläche einer Turbinenschaufel in der Impulsturbine. Bei einem Durchmesser von angenommenen 2cm ergibt sich, unter der Annahme, dass nur die äußere Hälfte der Fläche einer Schaufel angeströmt wird, eine effektive Fläche von 0,000157m² oder 157mm².When using a titanium alloy, strength values are usually between $300.0\frac{N}{m{m}^2}\text{and}1000.0\frac{N}{m{m}^2}.$

The area where the mass flow occurs corresponds to the area of a turbine blade in the impulse turbine. Assuming that only the outer half of the surface of a blade is exposed to the flow, an effective area of 0.000157m ² or 157mm ² results from an assumed diameter of 2cm.

als höchstmögliche Krafteinwirkung. Entscheidend ist die tatsächliche Krafteinwirkung im laufenden Betrieb.Taking into account the strength of such a turbine blade, the result is: $300.0\frac{N}{m{m}^2}\ast 157m{\mathrm{<figure-callout\; id="2"\; label="m"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">m</figure-callout>}}^2=47123N$

as the highest possible force. The decisive factor is the actual force applied during operation.

und ein Kolbenhub von 81 mm angeben. Dies ergibt eine „Strömungsgeschwindigkeiten“ von $\mathrm{8,37}\frac{m}{s}.$

Für einen Gasdrucklader, wie beispielsweise in einem G36-Gewehr Verwendung findet, treten Projektil-Beschleunigungen auf bis zu $\mathrm{920,0}\frac{m}{s}.$

In einer Glattrohrkanone von Rheinmetall, ebenfalls eine Waffe mit Gasdrucklader, werden bisweilen Geschwindigkeiten von $\mathrm{1640,0}\frac{m}{s}$

erreicht. (Quelle: http://www.whq-forum.de/cms/27.0.html; Unterpunkt: „Turm/Waffenanlage“; 20.10.2018, 17.00 Uhr) Aus diesen Überlegungen lässt sich schlussfolgern, dass in der Impulsturbine Strömungsgeschwindigkeiten von bis zu $\mathrm{1640,0}\frac{m}{s}$

und Drehzahlen von bis zu $\mathrm{19920,0}\frac{U}{min}$

realisierbar erscheinen.Another example of a high flow velocity is the gas pressure charger in a firearm or the accelerated piston in a car engine. In this case, expanding gases are actually used to accelerate other bodies - on the one hand the projectile, on the other hand the piston in the cylinder. For a motor vehicle cylinder, for example, a BMW M40 can have a speed of $6200\frac{U}{min}$

and specify a piston stroke of 81 mm. This results in a "flow velocity" of $8.37\frac{m}{s}.$

For a gas pressure loader, such as is used in a G36 rifle, projectile accelerations occur up to $920.0\frac{m}{s}.$

In a smooth barrel cannon from Rheinmetall, also a weapon with a gas pressure loader, speeds of $1640.0\frac{m}{s}$

reached. (Source: http://www.whq-forum.de/cms/27.0.html; sub-item: "Tower / weapon system"; October 20, 2018, 5:00 p.m.) From these considerations it can be concluded that in the impulse turbine flow velocities of up to to $1640.0\frac{m}{s}$

and speeds of up to $19920.0\frac{U}{min}$

appear feasible.

angenommen, während die Strömungsgeschwindigkeit maximal $1625\frac{m}{s}$

betragen soll. Daher lässt sich Folgendes annehmen: Sobald der Radius der Impulsturbine einen bestimmten Wert überschreitet, kann nicht mehr mit der höchstmöglichen Drehzahl operiert werden, ohne dass die höchstmögliche Geschwindigkeit überschritten wird. Sobald dieser bestimmte Radius der Impulsturbine unterschritten wird, kann nicht mehr mit der höchstmöglichen Geschwindigkeit operiert werden, ohne dass die höchstmögliche Drehzahl überschritten wird. Dieser Radius beträgt: $$r=\frac{{\nu }_I}{2\ast \pi \ast n}=\frac{1625m\ast s}{2\ast \pi \ast 332s}=\mathrm{0,779}m$$

Furthermore, the radius of the impulse turbine is of crucial importance. The speed of the turbine is at maximum $19220\frac{U}{min}=332.0\frac{1}{s}$

assumed while the flow velocity is maximum $1625\frac{m}{s}$

should be. The following can therefore be assumed: As soon as the radius of the impulse turbine exceeds a certain value, it is no longer possible to operate at the highest possible speed without exceeding the highest possible speed. As soon as the impulse turbine falls below this specific radius, it is no longer possible to operate at the highest possible speed without exceeding the highest possible speed. This radius is: $$r=\frac{{\nu }_{\mathrm{I.}}}{2\ast \pi \ast n}=\frac{1625m\ast s}{2\ast \pi \ast 332s}=0.779m$$

als höchstmöglicher Wert betrachtet. Zur Berechnung der Leistungsfähigkeit eines solchen Antriebssystems wird zunächst von der geforderten Leistung ausgegangen, die von der Impulsturbine erbracht werden muss. Die erforderliche Leistung wird dabei in Watt angegeben.Taking into account the assumption that the absolutely highest possible speed will most likely not be used during operation and that a downward regulation will take place slightly below this value, the following $n_{\mathrm{I.}}=320\frac{1}{s}$

considered the highest possible value. To calculate the performance of such a drive system, the required output that has to be provided by the impulse turbine is first assumed. The required power is given in watts.

$$P=2\ast \pi \ast n\ast M\ast {\eta }_{Turbine}\to M=\frac{P}{2\ast n\ast \pi \ast {\eta }_{Turbine}}$$

The following applies: $$\mathrm{P.}\left[\mathrm{P.}\mathrm{S.}\right]=\mathrm{P.}\ast 736\left[\mathrm{W.}\right]=\mathrm{P.}\ast 736\left[\frac{kG\ast m^2}{s^3}\right]$$

$$\mathrm{P.}=2\ast \pi \ast n\ast \mathrm{M.}\ast {\eta }_{Turbine}\to \mathrm{M.}=\frac{\mathrm{P.}}{2\ast n\ast \pi \ast {\eta }_{Turbine}}$$

It may be necessary to reduce the high speed of the impulse turbine to the maximum speed of the drive unit, namely the main rotor, propeller, etc. To ensure this, a spur gear must be used between the impulse turbine and the drive shaft. The efficiency of a multi-stage spur gear is usually around 90%. Combined with an impulse turbine, a final efficiency of 70-80% for the entire engine design can be assumed as an initial consideration. In this case, the efficiency of the gear must be included in the calculation of the Torque M flow into it. Torque M and speed n are further provided with the index I if they describe the speed or the torque of the impulse turbine. $$\mathrm{P.}=2\ast \pi \ast n\ast \mathrm{M.}\ast {\eta }_{Turbine}\ast {\eta }_{Getriebe}\to \mathrm{M.}=\frac{\mathrm{P.}}{2\ast n\ast \pi \ast {\eta }_{Turbine}\ast {\eta }_{Getriebe}}$$

The efficiency is considered to be the overall efficiency of the gearbox, which results from the efficiency of the gears, the bearings and that of the shaft, multiplied by the number of gear stages: $${\eta }_{Getriebe}={\left({\eta }_{welle}\ast {\eta }_{ZaHnra{\mathrm{<figure-callout\; id="2"\; label="d"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">d</figure-callout>}}^2}\ast {\eta }_{\mathrm{L.}aGe{r}^2}\right)}^{GesamtzaHlder\text{G}etriebestufen}$$

The individual efficiencies are assumed here with η = 0.9825. Modern transmissions can achieve values of up to η = 0.99 in each of their components. (Source: https://www.johannes-strommer.com/rechner/zugkraft-und-leistungs- punishment / efficiency / efficiency /; October 23, 2018, 6:10 p.m.) In an IRS engine, usually 2-speed or 3-speed engines -Step gears are used.

This results in a 2-stage gearbox: $${\eta }_{Getriebe}={\left(0.9825\ast {0.9825}^2\ast {0.9825}^2\right)}^2=0.83815$$

For a 3-stage gear this results: $${\eta }_{Getriebe}={\left(0.9825\ast {0.9825}^2\ast {0.9825}^2\right)}^3=0.76734$$

The gear ratio indicates how much the speed increases in each stage. The usual values are in the range from 3 to 5. The overall ratio is calculated as follows: $$i_{Ges}={\displaystyle \prod i_k}mit\left(k=GestriebestufenanzaHl\right)$$

The exact value of the individual gear ratios differs from engine to engine.

If the composition of the fuel-air mixture is known, an approximate statement about the fuel consumption of an engine can be made.

An aviation engine usually has a ratio of air mass flowing through to fuel weight consumed between 50: 1 and 200: 1.

A TV2-117 engine, for example, has an air throughput ṁ _air = 8.165 kg / s and a fuel _{consumption ṁ kerosene} = 454.772 kg / h at full power. (Source: http://www.jet-engine.net/miltsspec.html; 10/9/2018, 9:50 a.m.) $$\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{454.772\frac{kG}{H}}{8.165\frac{kG}{H}}=\frac{454.772\frac{kG}{H}}{8.165\ast 3600\frac{kG}{H}}=0.0155\frac{kG}{k{G}_{\mathrm{L.}uft}}$$

At full power, an Ivchenko AI-20D engine consumes ṁ _kerosene = 980.850 kg / h with an air throughput of ṁ _air = 20.865249 kg / s. (Source: http://www.jet-engine.net/civtfspec.html; October 23, 2018, 7:00 p.m.) $$\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{980.850\frac{kG}{H}}{20.865249\frac{kG}{H}}=\frac{980.850\frac{kG}{H}}{20.865249\ast 3600\frac{kG}{H}}=0.013058\frac{kG}{k{G}_{\mathrm{L.}uft}}$$

At full power, a Soloviev D-30KP engine has an air throughput ṁ _air = 290.0 kg / s and fuel _{consumption ṁ kerosene} = 5773.0 kg / h. (Source: http://www.jet-engine.net/civtfspec.html; October 23, 2018, 7:00 p.m.) $$\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{5773.0\frac{kG}{H}}{290.0\frac{kG}{H}}=\frac{5773.0\frac{kG}{H}}{290.0\ast 3600\frac{kG}{H}}=0.00553\frac{kG}{k{G}_{\mathrm{L.}uft}}$$

ist bei jedem Triebwerk unterschiedlich und kann auch über dem an dieser Stelle als höchsten Wert angegebenen $\frac{{\dot{m}}_{Kerosin}}{{\dot{m}}_{Luft}}=\mathrm{0,01550}$

liegen. Er lässt sich als Durchschnittwert nur schwer ermitteln. Es ist zur Beschreibung der Funktionsweise eines IRS-Triebwerks jedoch von keinem besonderen Belang, welche Form des Treibstoffs verwendet wird. Um im Weiteren die Rechnungen zu vereinfachen wird hier ein Wert $\frac{{\dot{m}}_{Kerosin}}{{\dot{m}}_{Luft}}=\mathrm{0,02}$

gewählt.The relationship $\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}$

is different for each engine and can also be higher than the highest value specified here $\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=0.01550$

lie. It is difficult to determine as an average value. However, it is of no particular importance which form of fuel is used in describing the operation of an IRS engine. In order to further simplify the calculations, a value is added here $\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=0.02$

elected.

nur für die Verwendung von Kerosin gültig sind. Würde an dieser Stelle ein anderer Treibstoff mit anderer Dichte, z.B. Bio-Kraftstoff oder Wasserstoff, verwendet, so würden sich die Werte für das Verhältnis $\frac{{\dot{m}}_{Treibstoff}}{{\dot{m}}_{Luft}}$

ändern.It should be noted that these values $\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}$

are only valid for the use of kerosene. If another fuel with a different density, e.g. bio-fuel or hydrogen, were used at this point, the values for the ratio would change $\frac{{\dot{m}}_{TreibstOff}}{{\dot{m}}_{\mathrm{L.}uft}}$

to change.

und der Heizwert mit $H_{Kerosin}=\mathrm{43,5}\frac{Mj}{kg}$

angegeben. $$E_{kin}=\frac{1}{2}\ast m\ast v_r{}^2\to v_r=\sqrt{\frac{2\ast E_{kin}}{m}}=\sqrt{\frac{2\ast 435000000\text{ kg}\ast {\text{m}}^2}{1kg\ast s^2}}$$

$$v_r=\mathrm{9237,379}\frac{m}{s}$$

For any indication of these values, the fuel used must be considered in terms of its density and its calorific value. The calorific value of a fuel can be represented as a speed v _T as an auxiliary variable. That means: If all of the energy stored in 1.0 kg of fuel were completely converted into kinetic energy, then this variable can be represented as speed v _T. In the case of kerosene, the density increases with ${\rho }_{KerOsin}=0.8\frac{\mathrm{<figure-callout\; id="3"\; label="k\; G\; m"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">k</figure-callout>}G}{m^{}}$

and the calorific value with $H_{KerOsin}=43.5\frac{\mathrm{M.}j}{kG}$

specified. $${\mathrm{E.}}_{kin}=\frac{1}{2}\ast m\ast {\mathrm{<figure-callout\; id="2"\; label="v\; r"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">v</figure-callout>}}_r{}^2\to v_r=\sqrt{\frac{2\ast {\mathrm{E.}}_{kin}}{m}}=\sqrt{\frac{2\ast 435000000\text{kg}\ast {\text{m}}^2}{1k\mathrm{<figure-callout\; id="2"\; label="G\; \ast \; s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">G</figure-callout>}\ast s^{}{}^2}}$$

$$v_r=\mathrm{9237,379}\frac{m}{s}$$

und der Heizwert $H_{H_2}=\mathrm{119,972}\frac{Mj}{m^3}.$

$$E_{kin}=\frac{1}{2}\ast m\ast v_r{}^2\to v_r=\sqrt{\frac{2\ast E_{kin}}{m}}=\sqrt{\frac{2\ast 119972000\text{ kg}\ast {\text{m}}^2}{1kg\ast s^2}}$$

$$v_r=\mathrm{15490,125}\frac{m}{s}$$

In the case of hydrogen, on the other hand, the density is ${\mathrm{<figure-callout\; id="2"\; label="\rho \; H"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">\rho </figure-callout>}}_{H_{}}=0.09\frac{\mathrm{<figure-callout\; id="2"\; label="k\; G\; m"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">k</figure-callout>}G}{m^{}}$

and the calorific value $H_{H_{}}$${{}_2}_{}=119.972\frac{\mathrm{<figure-callout\; id="3"\; label="M.\; j\; m"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">M.</figure-callout>}j}{m^{}}.$

$${\mathrm{E.}}_{kin}=\frac{1}{2}\ast m\ast {\mathrm{<figure-callout\; id="2"\; label="v\; r"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">v</figure-callout>}}_r{}^2\to v_r=\sqrt{\frac{2\ast {\mathrm{E.}}_{kin}}{m}}=\sqrt{\frac{2\ast 119972000\text{kg}\ast {\text{m}}^2}{1k\mathrm{<figure-callout\; id="2"\; label="G\; \ast \; s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">G</figure-callout>}\ast s^{}{}^2}}$$

$$v_r=15490.125\frac{m}{s}$$

verändern würde, müssen die beiden Geschwindigkeiten v_T in ein Verhältnis gesetzt werden. $$\frac{{\dot{m}}_{Treibstoff}}{{\dot{m}}_{Luft}}=\frac{{\dot{m}}_{Kerosin}}{{\dot{m}}_{Luft}}\ast \frac{v_{r,Kerosin}}{v_{r,Wasserstoff}}=\mathrm{0,02}\ast \frac{\mathrm{9237,379}}{\mathrm{15490,125}}=\mathrm{0,012}\frac{kg}{k{g}_{Luft}}$$

To determine how the factor is $\frac{{\dot{m}}_{TreibstOff}}{{\dot{m}}_{\mathrm{L.}uft}}$

would change, the two speeds v _T must be put in relation to each other. $$\frac{{\dot{m}}_{TreibstOff}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}\ast \frac{v_{r,KerOsin}}{v_{r,\mathrm{W.}asserstOff}}=0.02\ast \frac{\mathrm{9237,379}}{15490.125}=0.012\frac{kG}{k{G}_{\mathrm{L.}uft}}$$

verändern, wenn anstelle von Wasserstoff ein Kraftstoff mit anderer Dichte oder anderem Heizwert als Kerosin verwendet würde. $$\frac{{\dot{m}}_{Treibstoff}}{{\dot{m}}_{Luft}}=\frac{{\dot{m}}_{Kerosin}}{{\dot{m}}_{Luft}}\ast \frac{v_{r,Kerosin}}{v_{r,Kraftstoff}}$$

Similarly, the factor would be $\frac{{\dot{m}}_{TreibstOff}}{{\dot{m}}_{\mathrm{L.}uft}}$

change if a fuel with a different density or calorific value than kerosene were used instead of hydrogen. $$\frac{{\dot{m}}_{TreibstOff}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}\ast \frac{v_{r,KerOsin}}{v_{r,KraftstOff}}$$

These formulas describe the fact that the selected fuel has a considerable influence on the consumption of an engine. The individual subclasses of the IRS engine series are described in more detail in the following chapters.

Helicopters can better deal with the problem of poorly developed infrastructure than fixed-wing aircraft. Thanks to further technical development, helicopters can represent an alternative to fixed-wing aircraft, especially in regional and medium-haul transport. With the introduction of an IRS engine it is possible to counteract their high fuel consumption.

An IRS engine that is designed as a helicopter engine is referred to as an “impulse turbine helicopter engine”, or “IHT engine” for short. It is based on the idea of positioning a pulse turbine in the housing of the motor in such a way that it drives the main rotor via a shaft and, if necessary, via an additional spur gear. A spur gear unit differs from a bevel gear unit in that radial forces occur almost exclusively. When considering the efficiency of an IRS engine, including an IHT engine, a value of 70-80% can be assumed for the entire engine design.

In the schematic arrangement of the impulse turbine, jet engine and main rotor is shown. In the upper half of the picture, the two jet engines can be seen, which direct the mass flow, a fuel-air mixture, to the impulse turbine, shown as a "paddle wheel" in the middle of the picture. The arrows mark the direction of flow of the current. The impulse turbine is set in rotation by the mass flow, which is finally diverted through an exhaust gas opening. The main rotor of the helicopter is indicated by the rotor blades, which extend in a star shape from the center of the impulse turbine. It is located above the impulse turbine and both the hub of the impulse turbine and that of the main rotor are in one axis. However, the main rotor and the hub of the impulse turbine are not on one shaft. A spur gear is connected between them, which translates the speed of the impulse turbine into that of the main rotor.

In contrast to conventional helicopter drive units, the jet engines of the helicopter and the main rotor are not mechanically connected, the rotation of the main rotor takes place exclusively from the rotation of the impulse turbine. The other control components of the rotor, such as the blade adjustment or the tilting of the rotor, are not operated here. The tail rotor is also controlled separately from this arrangement.

The speed of the main rotor results from the regulation of the flow velocity of the mass flow and with the help of the translation through the spur gear. The performance of the impulse turbine depends on its speed, its radius and the mass of the fuel or fuel-air mixture flowing through it.

In order to estimate the performance of such a design, a helicopter model is used to check how the performance parameters change in relation to the current design of helicopter engines.

The engine power and the main rotor speed of the corresponding helicopter type are decisive for the calculation of an IHT engine.

These two values can then be used as output parameters for the calculation. Taking these output parameters into account, the entire calculation path back to the jet turbines, which generate the fuel mass flow, can be followed: $$\begin{array}{c}\underrightarrow{\underset{\mathrm{L.}eistunG\mathrm{P.};\mathrm{D.}reHzaHln}{{\displaystyle HauptrOtOr}}}\\ \underrightarrow{\underset{\mathrm{L.}eistunG\mathrm{P.};\mathrm{D.}reHzaHl{n}_{\mathrm{I.}}}{{\displaystyle \mathrm{S.}tirnradGetriebe}}}\\ \underrightarrow{\underset{Kraft\mathrm{F.};\mathrm{S.}trömunGsGescHwindiGkeit{v}_s}{{\displaystyle \mathrm{I.}mpulsturbine}}}\\ \underrightarrow{\underset{\mathrm{M.}assenstrOm{\dot{m}}_T}{{\displaystyle \mathrm{S.}traHlturbine}}}\\ TreibstOffverbraucH\end{array}$$

A Russian helicopter of the type "Mil Mi-8" is chosen as an example.

In most cases, a Mi-8 is powered by two TV2-117A engines, each delivering around 1670 hp. At full capacity, they consume a total of around 1140 liters / hour. (Source: http://www.jet-engine.net/miltsspec.html: October 2nd, 2018, 8:26 pm)

In this type of helicopter, the engines are mounted above the cockpit canopy. This arrangement is particularly well suited for the use of an IHT engine. In this case, the impulse turbine is located diagonally behind the red-covered exhaust opening on the side of the helicopter housing - in the turbine housing directly under the main rotor.

In order to estimate the expected fuel consumption, a pulse turbine with an output of 1670 hp is used. Assuming that two TV2-117A engines only generate 50% of 1670 HP as drive power (degree of reaction = 0.5; see p. 15), only one impulse turbine with an output of 1670 HP is required. $$1670\mathrm{P.}\mathrm{S.}=1229120\mathrm{W.}=1229120\frac{kG\ast m^2}{{\mathrm{<figure-callout\; id="3"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}^3}$$

angenommen. Mit Hilfe eines 3-Stufen-Getriebes wird diese dann auf die höchstmögliche des Hauptrotors reduziert. Für einen Hubschrauber vom Typ Mi-8 beträgt sie $n=185\frac{1}{min}.$

To determine the required force F, the required torque M _I , which is made up of the given power of 1670 hp and the speed of the turbine, is first used. The speed of the turbine is set to the maximum possible value of $n_l=320\frac{1}{s}$

accepted. With the help of a 3-stage gearbox, this is then reduced to the highest possible for the main rotor. For a Mi-8 helicopter it is $n=185\frac{1}{min}.$

In order to estimate the expected fuel consumption, a pulse turbine with an output of 1670 hp is used. $$\begin{array}{l}\mathrm{P.}=2\ast \pi \ast n_{\mathrm{I.}}\ast {\mathrm{M.}}_{\mathrm{I.}}\ast {\eta }_{Turbine}\ast {\eta }_{Getriebe}\to {\mathrm{M.}}_{\mathrm{I.}}=\frac{\mathrm{P.}}{2\ast n_{\mathrm{I.}}\ast {\pi }_{Turbine}\ast {\eta }_{Getriebe}}\\ \to {\mathrm{M.}}_{\mathrm{I.}}=\frac{1229120kG\ast {\mathrm{<figure-callout\; id="2"\; label="m"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">m</figure-callout>}}^2}{2\ast \pi \ast n_{\mathrm{I.}}\ast 0.90\ast 0.76734{\mathrm{<figure-callout\; id="3"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}^3}=\frac{283259.2298Nm}{n_{\mathrm{I.}}\ast 1s}\end{array}$$

The force F is the quotient of the torque M _I and the radius of the impulse turbine r. For the radius, a value can be obtained from various factors. The impulse turbine is limited in its diameter only by the dimensions of the surrounding housing and at the same time the largest possible radius is preferred. This means that in the optimal case the impulse turbine measures only a few centimeters or millimeters in diameter than the internal dimension of the housing. In this case, the dimensions of the engines can serve as a reference value for a Mi-8: Both engines are arranged next to one another and each have a diameter of 55 centimeters. In later models, the engines, type TV3-117, have a diameter of 66 cm. It can therefore be assumed that the engine casing measures approx. 132 cm in width. If a deviation of 6mm is assumed for the diameter of the impulse turbine, a radius r of 65.7cm can be specified.

At the beginning it was shown that in the case of impulse turbines with radii smaller than r = 0.779m, the highest possible speed of the impulse turbine can be used to calculate the fuel - this is also the case in this case. $$\frac{\mathrm{F.}}{n_{\mathrm{I.}}}=\frac{{\mathrm{M.}}_{\mathrm{I.}}}{r\ast n_{\mathrm{I.}}}=\frac{283259.2298Nm}{n_{\mathrm{I.}}\ast 0.657m\ast s}=\frac{\mathrm{431140,38022}N}{n_{\mathrm{I.}}\ast 1s}$$

dem Radius r = 0,657m und einem Wirkungsgrad von 90% und einem angeschlossenen 3-Stufen-Getriebe eine Leistung von 1670 PS zu verleihen, wird also eine Kraft von 1347,314 N benötigt.To a pulse turbine with the speed $n=320\frac{1}{s},$

To give the radius r = 0.657m and an efficiency of 90% and a connected 3-stage gearbox an output of 1670 HP, a force of 1347.314 N is required.

$$v_S=\mathrm{465,1519}\frac{m}{s}$$

This force results from the flow velocity v _{s of} the fuel-air mixture _{flow ṁ T.} This is to be equated with the exit speed of the mass flow from the jet engines. If a conventional jet engine such as a D-20P engine is used, which at full power has an air throughput of ṁ _air = 112.9445 kg / s and a fuel consumption of 3350.0 kg / h and generates a force of 52974 N in the process (Source: http://www.jet-engine.net/civtfspec.html; October 24, 2018, 1:32 p.m.), this speed v _{s is} approximately: $$v_{\mathrm{S.}}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub,\mathrm{D.}-\mathrm{20th}\mathrm{P.}}}{{\dot{m}}_{KerOsin,\mathrm{D.}-\mathrm{20th}\mathrm{P.}}+{\dot{m}}_{\mathrm{L.}uft,\mathrm{D.}-\mathrm{20th}\mathrm{P.}}}=\frac{52974N}{52974N\ast 0.000017555\frac{kG}{N\ast s}+112.9554\frac{kG}{s}}$$

$$v_{\mathrm{S.}}=465.1519\frac{m}{s}$$

The speed of the impulse turbine can be calculated from the flow velocity v _s , which is also the circumferential speed of the impulse turbine: $$n_{\mathrm{I.}}=\frac{v_{\mathrm{S.}}}{\left(2\ast \pi \ast r\right)}=\frac{465.1519\frac{m}{s}}{\left(2\ast \pi \ast 0.657m\right)}=112.6807\frac{1}{s}$$

The required force F is then: $$\mathrm{F.}=\frac{\mathrm{431140,38022}N}{112.6807\frac{1}{s}\ast 1s}=3826.2132N$$

It appears: $$52974N\ge 3826.2132N$$

A jet engine such as a D-20P engine provides sufficient power to drive an impulse turbine of the construction previously described. As a jet engine for driving this impulse turbine, an engine can therefore be used which corresponds to the performance parameters of a D-20P engine, but is structurally adapted in such a way that the exit / flow velocity v _{S is} maintained, while the force F generated is only 3826, 2132 N generated. This can be realized, for example, in the D-20P engine in an exact copy while maintaining all size and length ratios, but with a smaller dimension. By reducing the dimensions of the combustion chamber and the air inlet, the maximum volume of the engine is smaller and less thrust is generated than in the original engine. However, since all distance relationships within the engine are retained, the performance parameters do not change and the flow velocity v _S remains unchanged.

The mass flow of the fuel mixture ṁ _T can now be calculated from the force F determined in this way and the flow velocity v _S. $${\dot{m}}_T=\frac{\mathrm{F.}}{v_{\mathrm{S.}}}={\dot{m}}_{KerOsin}+{\dot{m}}_{\mathrm{L.}uft}={\dot{m}}_{\mathrm{L.}uft}\ast \left(\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}+1\right)$$

beträgt.The mass flow ṁ _T is made up of the mass of the air and that of the added fuel, with the ratio of both substances $\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}$

amounts to.

The fuel consumption ṁ _kerosene can then be determined from this result. $${\dot{m}}_{KerOsin}={\dot{m}}_T-{\dot{m}}_{\mathrm{L.}uft}={\dot{m}}_T\ast \frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{\mathrm{F.}}{v_{\mathrm{S.}}}\ast \frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}$$

impliziert, dass zum Erzielen eines möglichst geringen Treibstoffverbrauches ṁ_Treibstoff eine möglichst geringe Kraft F und/oder eine möglichst hohe Strömungsgeschwindigkeit v_S anzustreben ist.the equation $\frac{\mathrm{F.}}{v_{\mathrm{S.}}}={\dot{m}}_{KerOsin}+{\dot{m}}_{\mathrm{L.}uft}={\dot{m}}_{\mathrm{L.}uft}\ast \left(\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}+1\right)$

implies that in order to achieve the lowest possible fuel consumption ṁ _{fuel, the} lowest possible force F and / or _{the highest possible flow velocity v S should} be aimed for.

ebenfalls definiert.The force F is already given by the previous calculations. The fuel consumption in relation to the air mass flowing at the same time _{durch air} is given by the term $\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}$

also defined.

lässt sich nun der Treibstoffverbrauch abschätzen, den ein Mi-8 ṁ_Luft mit einem IHT-Triebwerk voraussichtlich erreichen wird: $${\dot{m}}_{Ker.}=\frac{F}{v_S}\ast \frac{{\dot{m}}_{Kerosin}}{{\dot{m}}_{Luft}}=\frac{\mathrm{3826,2132}N}{\mathrm{465,1519}\frac{m}{s}}\ast \mathrm{0,02}=\mathrm{0,165}\frac{kg}{s}=\mathrm{592,2525}\frac{kg}{h}\approx \mathrm{740,316}\frac{l}{h}$$

Taking into account the flow velocity v _s , the required force F and the $\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}$

you can now estimate the fuel consumption that a Mi-8 ṁ _Luft with an IHT engine is likely to achieve: $${\dot{m}}_{Ker.}=\frac{\mathrm{F.}}{v_{\mathrm{S.}}}\ast \frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{3826.2132N}{465.1519\frac{m}{s}}\ast 0.02=0.165\frac{kG}{s}=592.2525\frac{kG}{H}\approx \mathrm{740,316}\frac{l}{H}$$

The conclusion can be drawn: With the help of a pulse turbine construction, the fuel consumption of a conventional helicopter model such as the Mi-8 can be reduced to approx. 70% of the original value.

Wird andererseits ein ineffizienteres Triebwerkspaar verwendet, was eine geringere Strömungsgeschwindigkeit erzeugt $\left(\text{z}\text{.B }\mathrm{420,0}\frac{m}{s}\right),$

so erhöht sich der Treibstoffverbrauch unter Beibehaltung aller anderen Parameter in diesem Fall auf ca. $\mathrm{908,046}\frac{l}{h}.$

Auch dieser Wert ist allerdings vergleichsweise niedrig verglichen mit dem gegenwertigen Verbrauch von ca. $1100\frac{l}{h}.$

It should be noted that all values used are based on rough estimates and are heavily dependent on the components actually used. This can increase fuel consumption in real terms. If a further gear stage is added to the spur gear in the calculation described above, the fuel consumption increases to approx. $808.64\frac{l}{H}.$

On the other hand, it becomes a more inefficient one Engine pair used, which creates a lower flow velocity $\left(\text{z}\text{.B}420.0\frac{m}{s}\right),$

in this way, the fuel consumption increases to approx. $\mathrm{908,046}\frac{l}{H}.$

However, this value is also comparatively low compared to the current consumption of approx. $1100\frac{l}{H}.$

The various IRS engine designs, namely the IHT engines, TPI and the other IRS engine designs, aim to achieve an optimal increase in the efficiency of the entire drive, so the use of machine elements with high levels of efficiency can be assumed.

An IHT engine can theoretically also be installed in a helicopter, in which the arrangement of the engines is not as optimal as in a Mi-8. An example of this is a CH-53G helicopter that is currently in service with the German Federal Armed Forces.

With this pattern it is too clear to see that the engines are attached to the side of the fuselage and are not a straight extension to the casing that is supposed to contain the impulse turbine. In this case, the use of conventional jet engines is relatively inexpedient. It is possible to use small special injection nozzles that inject the fuel as well as direct air into the engines via intake manifolds. This then results in a fuel-air mixture, which is finally fed into the central housing to the impulse turbine via feed lines.

The CH-53G, as it is in service with the German Armed Forces, is a license production of the American Sikorsky S-65. The engines are manufactured by MTU and designated as T64-MTU-7. With take-off power, they each produce 3975 hp and consume approx. 1000 liters per hour.

In order to implement an IHT engine in this helicopter model, considerations regarding the flow velocity of the mass flow and the power achieved are of primary importance.

The engine parameters such as the radius r of the impulse turbine and its speed result from the structural conditions.

The radius r can be estimated from the dimensions of the engine housing. The entire width of the fuselage cell without engines is approx. 2.80 m, therefore a width of 1.50 m can be assumed for the width of the turbine housing. An assumption of 1.25 m is therefore made for the diameter of the impulse turbine and a value of 0.625 m ultimately results for its radius r.

• Ein Strahltriebwerk konventionellen Typs lässt sich in einem CH-53G-Hubschrauber schlecht verwenden, anstelle dessen werden spezielle Einspritzdüsen verwendet.

For the flow velocity v _S , the flow velocity of the fuel flow, only estimates are also possible:

• A conventional type of jet engine is difficult to use in a CH-53G helicopter and special fuel injectors are used instead.

bewegen. Als wahrscheinlich höchstmögliche Strömungsgeschwindigkeit wurde eingangs $\mathrm{1625,0}\frac{m}{s}$

definiert und als höchstmögliche Drehzahl $n_I=320\frac{1}{s}$

festgelegt. Da beide Werte von dem des Radius' abhängen und dieser geringer als 0,779m (s. S. 17) ist, wird für die weitere Berechnung die höchstmögliche Drehzahl $n_I=320\frac{1}{s}$

als Basiswert verwendet. Die Hauptrotor-Drehzahl beträgt wie auch bei der $\text{Mi}-8n=185\frac{1}{min}=\mathrm{3,083}\frac{1}{s}.$

Therefore, in this case, a higher flow velocity is assumed than in the case of a jet engine, in which the values of the velocity v _{S are} usually between $300.0\frac{m}{s}\text{and}500.0\frac{m}{s}$

move. At the beginning, the probably highest possible flow velocity was stated $1625.0\frac{m}{s}$

defined and as the highest possible speed $n_{\mathrm{I.}}=320\frac{1}{s}$

set. Since both values depend on that of the radius and this is less than 0.779m (see p. 17), the highest possible speed is used for the further calculation $n_{\mathrm{I.}}=320\frac{1}{s}$

used as base value. The main rotor speed is the same as with the $\text{Wed}-\mathrm{8th}n=185\frac{1}{min}=3.083\frac{1}{s}.$

auf $\mathrm{3,083}\frac{1}{s}$

kann hier von einer Verwendung eines 3-Stufen-Getriebes ausgegangen werden. Der Wirkungsgrad dieses Getriebes wird bei η_Getriebe = 0,76734 angesetzt. $$P=3975PS=2925600W=2925600\frac{kg\ast m^2}{s^3}$$

$$M_I=\frac{2925600kg\ast m^2}{2\ast \pi \ast 320\frac{1}{s}\ast \mathrm{0,90}\ast \mathrm{0,76734}{s}^3}=\mathrm{2106,9526}Nm$$

Due to the high gear ratio of $320.0\frac{1}{s}$

on $3.083\frac{1}{s}$

a 3-stage gearbox can be assumed here. The efficiency of this gear is set at η _gear = 0.76734. $$\mathrm{P.}=3975\mathrm{P.}\mathrm{S.}=2925600\mathrm{W.}=2925600\frac{kG\ast m^2}{{\mathrm{<figure-callout\; id="3"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}^3}$$

$${\mathrm{M.}}_{\mathrm{I.}}=\frac{2925600kG\ast {\mathrm{<figure-callout\; id="2"\; label="m"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">m</figure-callout>}}^2}{2\ast \pi \ast 320\frac{1}{s}\ast 0.90\ast 0.76734{\mathrm{<figure-callout\; id="3"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}^3}=2106.9526Nm$$

The gear ratio indicates how much the speed increases in each stage. The usual values are in the range from 3 to 5. The overall ratio is calculated as follows: $$i_{Ges}={\displaystyle \prod i_k}mitk=GetriebestufenanzaHl$$

The following applies in this case: $$i_{Ges}={\displaystyle \prod i_3={\mathrm{<figure-callout\; id="1"\; label="i"\; filenames="DE102019130861B4\_0150.png"\; state="\{\{state\}\}">i</figure-callout>}}_1\ast i_2\ast {\mathrm{<figure-callout\; id="3"\; label="i"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">i</figure-callout>}}_3}$$

For example, if a value of i _k ≈ 4.7 is used for each stage, so an overall translation results from i _ges = 103.823. For the speed of the impulse turbine this means: n _I = i _total * n = 103.823 * 3.083 ≈ 320.0 _. This means that a 3-stage gearbox is sufficient for realizing this IHT engine.

The following applies to the flow velocity v _S in the impulse turbine: $$v_s=2\ast \pi \ast n_{\mathrm{I.}}\ast r=2\ast \pi \ast 320.0\frac{1}{s}\ast 0.625m=1256.637\frac{m}{s}$$

The force required is as follows: $$\mathrm{F.}=\frac{{\mathrm{M.}}_{\mathrm{I.}}}{r}=\frac{2106.9526Nm}{0.625m}=3371.14016N$$

im laufenden Betrieb. Ein Einsatz einer solchen Impulsturbine kann daher als praktikabel angesehen werden.The maximum permissible force acting on a turbine blade is 47123 N (see p. 17) and is thus far above the actual force acting at a flow velocity v _S of $1256.637\frac{m}{s}$

in operation. The use of such a pulse turbine can therefore be regarded as practicable.

For the fuel consumption then follows: $${\dot{m}}_{KerOsin}=\frac{\mathrm{F.}}{v_{\mathrm{S.}}}\ast \frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{3371.14016N}{1256.637\frac{m}{s}}\ast 0.02=0.0537\frac{kG}{s}=193.153\frac{kG}{H}\sim 241.5\frac{l}{H}$$

The fuel consumption can be reduced noticeably to below 50% under given circumstances.

führen zu einem Verbrauch von ca. $\mathrm{645,0}\frac{l}{h}.$

Dieser Wert ist ebenfalls unterhalb des ursprünglichen von $\mathrm{1000,0}\frac{l}{h}$

zu verzeichnen.Comparative considerations with, for example, a lower flow velocity of only $\mathrm{470,000}\frac{m}{s}$

lead to a consumption of approx. $645.0\frac{l}{H}.$

This value is also below the original of $1000.0\frac{l}{H}$

to be recorded.

Since fixed-wing aircraft work more economically than comparable helicopters, their importance will most likely increase in the future, especially in long-haul transport across the region and in long-term special missions. Statistically, at the present time the group of jet engines is most efficient in high-speed operation, while aircraft powered by piston engines such as a Cessna C172 are best suited for the lowest speed range. The speed range of turboprop machines like the A400M is relatively midway between that of jet engines and that of piston engines. In the case of turboprop engines, a bypass ratio is very seldom used, since the propeller has no casing and the air mass flow flowing through the propeller is therefore not directed through the engine. In this context, the “bypass ratio” in turboprop engines describes the ratio of the propeller air mass flow and the thrust provided by the core flow. This ratio is typically around 10/1 to sometimes 40/1. This fact points to the relatively low fuel consumption of turboprop engines. At the same time, most of the turboprop engines currently in use are only suitable for lower altitudes than jet engines. This is due to the fact that they have a higher air throughput in order to achieve their full thrust and are therefore dependent on higher air pressure. Furthermore, in most cases they do not achieve the same thrust performance as jet engines of comparable size, which consequently limits the flight performance of most turboprop aircraft.

The most efficient turboprop engines in history are the Soviet MK-12MV, which is still in service with the Ukrainian Antonov An-22, and the EPI TP400, which powers the A400M. Both engines enable flight performances that are comparable to jet-powered medium-haul aircraft. On record flights, other turboprop aircraft also reached altitudes of 13,000m, which is higher than the maximum altitude of many jet-powered medium-haul aircraft with a comparable weight and range. Turboprop engines can therefore be considered very well for a design that can operate in both the upper and lower speed ranges.

In order to further optimize these engines, the idea of a pulse turbine can also be used here. Such a construction is called a “turboprop impulse turbine engine”, or “TPI engine” for short.

The mode of operation of such a TPI engine is linked to the description of the IHT engine: a pulse turbine is installed in the interior of the turbine housing in a vertical orientation, driven by a mass flow and finally actuates the propeller via a multi-stage gearbox.

In the a TPI engine is shown from the frontal perspective. In the center is the impulse turbine, which is driven by a fuel mass flow that is introduced via a feed line at the upper edge. This is where the injection nozzle is located, which feeds the fuel-air mixture into the turbine. The air mass is sucked in via a special ramjet engine or a special turbine.

A special feature of a TPI engine results from the fact that the air flow and the axis of rotation of the impulse turbine are aligned. The plane of rotation of the impulse turbine is therefore perpendicular to the flight direction of the aircraft. For the sucked-in air mass, this means that it or the fuel-air mixture is guided into the impulse turbine in an arc, deflected by 90 °. As a result, friction losses can occur in the mass flow, which do not occur with a straight course as in an IHT engine. With an appropriate choice of material and wide arcuate ducting, however, these losses are very limited.

Compared to currently used turboprop engines, no core flow is generated in a TPI engine. As in an airplane with a piston engine, the propulsive power comes exclusively from the power of the propeller. In a TPI engine, therefore, the propulsion force component that results from the core flow is omitted. This would result in a ratio of approx. 10/1 to partially 40/1 Propulsion losses of 2.5 - 10%. However, this can be compensated for by a higher performance of a TPI engine and a higher speed of the propeller or longer propeller blades.

The high speeds result from the fact that the radii of TPI turbines are usually smaller than those of IHT engines, while the required power is usually higher.

A TP400 should serve as an example of a possible application. This is one of the most powerful turboprop engines today and is used on the Airbus A400M, a military transport aircraft.

The air inlet of this engine can be seen on the lower side. In the example with a TPI engine, a ramjet engine is connected, which directs the air mass, which is finally mixed with the fuel, into the impulse turbine.

The diameter of an EPI TP400 is around 0.92 m. For a possible radius r of the impulse turbine, with a deduction of a few centimeters, this results in about r = 0.452 m. The power of a TP400 is estimated at 8000 HP in cruise flight, comparable to a Soviet NK-12MV engine. At an altitude of 36,100 feet, this makes 8000PS and, overall, is slightly above that of the TP400 in terms of performance. $$8000\mathrm{P.}\mathrm{S.}=5888000\mathrm{W.}=5888000\frac{kG\ast m^2}{{\mathrm{<figure-callout\; id="3"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}^3}$$

(Quelle: https://defence.pk/pdf/threads/airbus-a400-m-detailed-analysis-including-pilot-reports.273682/; 20.10.2018, 15.45 Uhr) Für jedes Triebwerk bedeutet dies einen Verbrauch von $\mathrm{1091,5}\frac{l}{h}.$

When cruising, an A400M with standard TP400 engines consumes about $4366\frac{l}{H}.$

(Source: https://defence.pk/pdf/threads/airbus-a400-m-detailed-analysis-including-pilot-reports.273682/; October 20, 2018, 3:45 p.m.) This means a consumption of $1091.5\frac{l}{H}.$

With the help of the given data, the fuel consumption can now be estimated if a TPI engine is used at this point. First, the required torque is calculated for this. The highest possible speed is used for the flow velocity in the impulse turbine: $$v_s=2\ast \pi \ast n_{\mathrm{I.}}\ast r=2\ast \pi \ast 320.0\frac{1}{s}\ast 0.452m=908.799\frac{m}{s}$$

(Quelle: https://www.easa.europa.eu/sites/default/files/dfu/TCDS E033 Issue%2006 1.0 0.pdf; 20.10.2018; 18.00 Uhr) Für das Übersetzungsverhältnis bedeutet dies bei einer Übersetzung von 4,71405 in jeder Stufe: $$i_{ges}={\displaystyle \prod i_2=i_1\ast i_2=\mathrm{4,71405}\ast \mathrm{4,71405}\approx \mathrm{22,}\overline{2}}$$

$$n_I=n_{Propeller}\ast i_{ges}=\mathrm{864,0}\frac{1}{60s}\ast \mathrm{22,}\overline{2}=\mathrm{320,0}\frac{1}{s}$$

In contrast to an IHT engine, a gearbox with only 2 stages can be used in a TPI engine, as the necessary ratio of impulse turbine speed to propeller speed is not as high as in an IHT engine. With a TP400 the maximum speed of the propeller is included $864.0\frac{U}{min}.$

(Source: https://www.easa.europa.eu/sites/default/files/dfu/TCDS E033 Issue% 2006 1.0 0.pdf; 10/20/2018; 6:00 p.m.) For the translation ratio, this means with a translation of 4 , 71405 in each level: $$i_{Ges}={\displaystyle \prod i_2={\mathrm{<figure-callout\; id="1"\; label="i"\; filenames="DE102019130861B4\_0150.png"\; state="\{\{state\}\}">i</figure-callout>}}_1\ast i_2=4.71405\ast 4.71405\approx 22\overline{2}}$$

$$n_{\mathrm{I.}}=n_{\mathrm{P.}rOpeller}\ast i_{Ges}=864.0\frac{1}{60s}\ast 22\overline{2}=320.0\frac{1}{s}$$

$${\mathrm{\eta }}_{Getriebe}={\left(\mathrm{0,9825}\ast {\mathrm{0,9825}}^2\ast {\mathrm{0,9825}}^2\right)}^2=\mathrm{0,83815}$$

A 2-stage gearbox is therefore sufficient to convert the speed of the impulse turbine into that of the propeller. The following applies to the efficiency of the gear unit: $${\mathrm{\eta }}_{Getriebe}={\left({\mathrm{\eta }}_{welle}\ast {\mathrm{\eta }}_{ZaHnra\mathrm{<figure-callout\; id="2"\; label="d"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">d</figure-callout>}}{}^2\ast {\mathrm{\eta }}_{\mathrm{L.}aGer}{}^2\right)}^{GesamtzaHlderGetriebestufen}$$

$${\mathrm{\eta }}_{Getriebe}={\left(0.9825\ast {0.9825}^2\ast {0.9825}^2\right)}^2=0.83815$$

$$M_I=\mathrm{3882,1624}Nm$$

For the torque it finally follows: $${\mathrm{M.}}_{\mathrm{I.}}=\frac{\mathrm{P.}}{2\ast n_{\mathrm{I.}}\ast \pi \ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}=\frac{5888000kG\ast {\mathrm{<figure-callout\; id="2"\; label="m"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">m</figure-callout>}}^2\ast \mathrm{<figure-callout\; id="2"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}{2\ast \pi \ast 320.0{\mathrm{<figure-callout\; id="3"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}^3\ast 0.90\ast 0.83815}$$

$${\mathrm{M.}}_{\mathrm{I.}}=3882.1624Nm$$

The force required is as follows: $$\mathrm{F.}=\frac{{\mathrm{M.}}_{\mathrm{I.}}}{r}=\frac{3882.1624Nm}{0.452m}=8588.855N$$

For the fuel consumption then follows: $${\dot{m}}_{KerOsin}=\frac{\mathrm{F.}}{v_s}\ast \frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{8588.855N}{908.799\frac{m}{s}}\ast 0.02=0.189\frac{kG}{s}=680.456\frac{kG}{H}\sim 850.57\frac{l}{H}$$

eine Reduktion um 22% des ursprünglichen Wertes dar. Dies ist zum Teil auch der Tatsache geschuldet, dass es sich bei dem TP400 um ein sehr leistungsstarkes Triebwerk handelt - ein Triebwerk mit niedrigem Leistungsgewicht. „Leistungsgewicht“ beschreibt den Quotienten aus Masse und Leistung eines Antriebs, Fortbewegungsmittels, Kraftwerks oder auch bei Mensch oder Tier. Im Falle eines TP400 Triebwerks bedeutet das, dass eine verhältnismäßig hohe Leistung einem „kleinen“ Triebwerk abgefordert wird. „Kleines“ Triebwerk bedeutet, wie gezeigt, kleiner Radius in der Impulsturbine und daher auch eine geringe Strömungsgeschwindigkeit.The savings cannot be compared to those of IHT engines, but they are remarkably high. In the example shown, the $850.57\frac{l}{H}$

a reduction of 22% of the original value. This is partly due to the fact that the TP400 is a very powerful engine - an engine with a low power-to-weight ratio. "Power-to-weight ratio" describes the quotient of the mass and power of a drive, means of transport, power plant or even for humans or animals. In the case of a TP400 engine, this means that a relatively high output is required from a "small" engine. “Small” engine means, as shown, a small radius in the impulse turbine and therefore also a low flow velocity.

Another example, such as a Ukrainian AI-20 engine, produces less power in a larger engine.

An AI-20 is a Ukrainian, formerly Soviet, engine that was developed in the 1950s. It is still used today on aircraft such as the Antonov An-12.

On the lower side of the engines, the air intake ports can be seen, which direct the air flow into the interior of the turbines. In the case of a TPI engine, the impulse turbine would be located at this point.

(Quelle: http://www.motorsich.com/eng/products/aircraft/tr/ai-20/; 20.10.2018, 20.20 Uhr) Die Breite des Triebwerksgehäuses beträgt nach Herstellerangaben 0,842m. Das bedeutet für die Impulsturbine einen Radius r = 0,400m. Zur Berechnung des Treibstoffverbrauchs in dieser Version des TPI-Triebwerks wird wieder eine Drehzahl in der Impulsturbine $n_I=19200\frac{U}{min}=320\frac{1}{s}$

min s sowie ein Getriebewirkungsgrad von η_Getreibe= 0,83815 angenommen. Für das Drehmoment folgt schließlich: $$M_I=\frac{P}{2\ast n_I\ast \pi \ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}=\frac{1832640kg\ast m^2\ast s}{2\ast \pi \ast \mathrm{320,0}{s}^3\ast \mathrm{0,90}\ast \mathrm{0,83815}}$$

$$M_I=\mathrm{1208,32304}Nm$$

Such an engine makes about 2490PS (= 1832640W) when cruising and consumes it $653.625\frac{l}{H}$

(Source: http://www.motorsich.com/eng/products/aircraft/tr/ai-20/; October 20, 2018, 8:20 p.m.) According to the manufacturer, the width of the engine housing is 0.842 m. For the impulse turbine this means a radius r = 0.400m. To calculate the fuel consumption in this version of the TPI engine, a speed is again used in the impulse turbine $n_{\mathrm{I.}}=19200\frac{U}{min}=320\frac{1}{s}$

min s and a gear _{efficiency of η gear units} = 0.83815 assumed. For the torque it finally follows: $${\mathrm{M.}}_{\mathrm{I.}}=\frac{\mathrm{P.}}{2\ast n_{\mathrm{I.}}\ast \pi \ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}=\frac{1832640kG\ast {\mathrm{<figure-callout\; id="2"\; label="m"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">m</figure-callout>}}^2\ast \mathrm{<figure-callout\; id="2"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}{2\ast \pi \ast 320.0{\mathrm{<figure-callout\; id="3"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}^3\ast 0.90\ast 0.83815}$$

$${\mathrm{M.}}_{\mathrm{I.}}=\mathrm{1208,32304}Nm$$

The force required is as follows: $$\mathrm{F.}=\frac{{\mathrm{M.}}_{\mathrm{I.}}}{r}=\frac{\mathrm{1208,32304}Nm}{0.400m}\mathrm{3020,8076}N$$

The flow velocity v _s : $$v_s=2\ast \pi \ast n_{\mathrm{I.}}\ast r=2\ast \pi \ast 320.0\frac{1}{s}\ast 0.400m=\mathrm{804,247719}\frac{m}{s}$$

gilt dann für den Treibstoffverbrauch: $${\dot{m}}_{Ker.}=\frac{F}{v_s}\ast \frac{{\dot{m}}_{Kerosin}}{{\dot{m}}_{Luft}}=\frac{\mathrm{3020,8076}N}{\mathrm{804,247719}\frac{m}{s}}\ast \mathrm{0,02}=\mathrm{0,075121}\frac{kg}{s}=\mathrm{270,44}\frac{kg}{h}\sim \mathrm{338,046}\frac{l}{h}$$

Taking into account the factor $\frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=0.02$

then applies to fuel consumption: $${\dot{m}}_{Ker.}=\frac{\mathrm{F.}}{v_s}\ast \frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{\mathrm{3020,8076}N}{\mathrm{804,247719}\frac{m}{s}}\ast 0.02=0.075121\frac{kG}{s}=270.44\frac{kG}{H}\sim \mathrm{338,046}\frac{l}{H}$$

This value has a ratio of almost 1: 2 to the original value. With a TPI engine, an An-12 would only use 52% of the fuel to cover its current performance range.

Jet engines have been established as the most widely used propulsion system in aviation since the last quarter of the 20th century. They have largely replaced turboprop engines in both the civil and military sectors. They are characterized by a high level of economic efficiency in the upper speed range from Mach 0.6 and can usually also be used in higher air layers than turboprop engines. Your role in aviation will therefore be of central importance in the future. With a few exceptions, jet engines that operate in the subsonic range are designed in such a way that part of the air mass sucked in is bypassed by the engine core. These thrusters are known as dual-flow or bypass thrusters. Logically, the bypassed air mass is referred to as “sheath flow” or “bypass flow”. The air mass that is directed through the engine core is called the “core flow”. The air mass in the core and secondary flow is accelerated by a compression stage and, when it exits, generates a backward force that drives the aircraft according to the recoil principle. In the core flow, the air mass is heated and accelerated together with the ignited fuel; in the sheath flow, only the air mass is compressed and accelerated. The turbine in the engine core also drives the turbine blades in the low pressure range - those that are in the front of the engine. Nowadays, high bypass ratios are often sought in the design of jet engines. Comparable to turbo-prop engines, where the "ratio" is greater than 10/1, with high bypass flow ratios this means that a high proportion of the thrust generated comes from compressed, accelerated air. The higher this proportion, the lower the proportion of the fuel-permeated core flow and, consequently, the fuel consumption is also low - lower than with a single-flow engine, where the entire air mass is passed through the engine as a core flow and the entire thrust comes from one fuel. Air mixture is generated. Modern engines like the Rolls-Royce Trent XWB achieve bypass ratios of up to 9.6 / 1.

IRS engines that are designed as jet engines are referred to as "jet turbine impulse engines" - "STI engines" for short. They have the peculiarity that they do not have a core stream. The thrust of the engine occurs only through the acceleration in the "bypass flow". The impulse turbine that drives the fan is located in the engine core. The plane of rotation is perpendicular to the flow direction of the air masses in the engine and thus also perpendicular to the propulsion direction of the aircraft. Therefore, as in the TPI engine, the mass flow in the impulse turbine and the air mass flow in the engine casing are hermetically shielded from one another. A ramjet engine is used to generate the mass flow in the impulse turbine, through which air is sucked in, accelerated and then fed into the turbine mixed with fuel. The intake manifold for this air flow may be in the hub of the engine or at another convenient location on the aircraft.

In the sketch of an STI engine is shown from the frontal perspective. In the center of the picture the impulse turbine is shown, which is set in rotation by the mass flow supplied above. over the low-pressure turbine is driven by a shaft and, if necessary, an interposed gearbox - its blades are shown in the form of a drawing as a wreath around the impulse turbine. The area behind the low-pressure turbine represents the jacket chamber of the engine. In it, the air mass passed through is accelerated and expelled as propulsion-generating force on the rear side of the engine. The fuel-air mixture mass flow that drives the impulse turbine in the engine core is hermetically shielded from the air mass that is passed through the engine casing chamber. It is possible that the exhaust air of the "core flow" is expelled from the engine as an additional mass flow and provides additional thrust. However, this is not taken into account when calculating the engine performance, as the final design of such an STI engine has not been fully clarified. Assuming that the “core flow” is not available as thrust, the functionality of the STI engine is comparable to that of a TPI engine: The thrust of the entire engine then comes exclusively from the air mass that is passed through the jacket chamber of the engine becomes. To calculate the performance of such an STI engine, the flow velocity of the bypass flow and the thrust generated by the engine are important. Unlike helicopter and turboprop engines, jet engines are usually defined by their thrust.

When designing an STI engine, it must be noted that the ratio of the radius of the impulse turbine to that of the low-pressure turbine is decisive for engine performance and fuel consumption. The larger the low-pressure turbine, the more air mass can be moved through the engine while maintaining the same flow rate, and the more thrust it generates. On the other hand, if the radius of the entire engine construction does not increase, this means that the radius of the impulse turbine in the engine core is reduced. However, a smaller impulse turbine radius leads to an increased energy requirement and thus to an increased fuel requirement while maintaining the same output. The ratio of the two radii to each other must therefore be determined separately for each engine. It depends on the engine dimensions and the desired thrust.

erhöhen. Um eine bessere Vergleichbarkeit mit herkömmlichen kommerziell eingesetzten Triebwerken zu erzielen, genügt es an diesem Punkt, die Austrittsgeschwindigkeit des Luftstroms aus dem Triebwerk auf einen vergleichbaren Wert zu „erhöhen“.In general, it can be stated that a high flow velocity in the engine casing leads to a low air mass requirement to generate a certain thrust. Thanks to the conical shape of the engine, the flow speed can be increased almost continuously to a certain value. Theoretically, this value could be based on that of conventional rocket engines of about $3000.0\frac{m}{s}$

increase. In order to achieve a better comparison with conventional commercially used engines, it is sufficient at this point to "increase" the exit speed of the air flow from the engine to a comparable value.

An example with a relatively high flow rate of the air mass is a Soviet D-20P engine (see p. 20 ff). Its flow rate, at the same time its exit velocity, is $v_{\mathrm{A.}}=465.1519\frac{m}{s}.$

$$v_A=\mathrm{361,482}\frac{m}{s}$$

Another example is an IAE V2533-A5 engine that is now used on an Airbus A321. (Source: http://www.jet-engine.net/civtfspec.html: October 24, 2018, 8:30 p.m.) Its exit speed is: $$v_{\mathrm{A.}}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub,V2533}}{{\dot{m}}_{KerOsin,V2533}+{\dot{m}}_{\mathrm{L.}uft,V2533}}=\frac{139721.87222N}{139721.87222N*0.0000102\frac{kG}{N\ast s}+385.0999\frac{kG}{s}}$$

$$v_{\mathrm{A.}}=\mathrm{361,482}\frac{m}{s}$$

In the same way, the flow velocities can also be calculated for other engines. (e.g. Solovyev D-30-1 → $v_{\mathrm{A.}}=\mathrm{525,077}\frac{m}{s})$

verwendet. Die Austrittsgeschwindigkeit multipliziert mit der Schubkraft ergibt die Leistung des Triebwerks. Dieses Produkt lässt sich in die Formel zur Leistungsberechnung einsetzen (s. S. 17): $$P\left[PS\right]=P\ast 736\left[W\right]=P\ast 736\left[\frac{kg\ast m^2}{s^3}\right]=F_{Schub}\ast v_{Durchfluss}\left[\frac{kg\ast m^2}{s^3}\right]$$

In view of the fact that there is a different optimal ratio of these two radii for each engine, which also has an influence on the flow rate, a speed is used here as an example $v_{\mathrm{A.}}=462.5\frac{m}{s}$

used. The exit speed multiplied by the thrust gives the power of the engine. This product can be used in the formula for power calculation (see p. 17): $$\mathrm{P.}\left[\mathrm{P.}\mathrm{S.}\right]=\mathrm{P.}\ast 736\left[\mathrm{W.}\right]=\mathrm{P.}\ast 736\left[\frac{kG\ast m^2}{s^3}\right]={\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHfluss}\left[\frac{kG\ast m^2}{s^3}\right]$$

angepasst werden.The following should be noted for the velocity v _flow : It differs from engine to engine and depends on the thrust and the air throughput of the respective engine. The force F actually required, which must be generated by the impulse turbine, must therefore with the factor $\frac{v_{\mathrm{D.}urcHfluss}}{v_{\mathrm{A.}}}$

be adjusted.

als Basiswert verwendet. Ist der Radius geringer als 0,779m, so wird zur Berechnung die Drehzahl $n_I=320\frac{1}{s}$

als Basiswert verwendet.The length of the radius of the impulse turbine determines which formula is most appropriate to use to calculate fuel consumption (see p. 17). If it is greater than 0.779m, the flow velocity is used for the calculation $v_{\mathrm{S.}}=1625\frac{m}{s}$

used as base value. If the radius is less than 0.779m, the speed is used for the calculation $n_{\mathrm{I.}}=320\frac{1}{s}$

used as base value.

$$F_{Schub}\ast v_A\ast \frac{v_{Durchfluss}}{v_A}=2\ast \pi \ast n_I\ast M_I\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}$$

$$M_I=\frac{F_{Schub}\ast v_A\ast v_{Durchfluss}}{2\ast n_I\ast \pi \ast v_A\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}$$

$$Strmungsgeshwindigkeit:v_S=2\ast \pi \ast n_I\ast r$$

$$M_I=\frac{F_{Schub}\ast v_A\ast v_{Durchfluss}}{\frac{v_S}{r}\ast v_A\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}$$

$$F=\frac{M_I}{r}=\frac{F_{Schub}\ast v_A\ast v_{Durchfluss}}{r\ast \frac{v_S}{r}\ast v_A\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}=\frac{F_{Schub}\ast v_A\ast v_{Durchfluss}}{v_S\ast v_A\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}$$

First, the calculation takes place with the aid of the flow velocity v _s : $$\mathrm{P.}\left[\mathrm{P.}\mathrm{S.}\right]=\mathrm{P.}\ast 736\left[\mathrm{W.}\right]=\mathrm{P.}\ast 736\left[\frac{kG\ast m^2}{s^3}\right]={\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHfluss}\left[\frac{kG\ast m^2}{s^3}\right]$$

$${\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{A.}}\ast \frac{v_{\mathrm{D.}urcHfluss}}{v_{\mathrm{A.}}}=2\ast \pi \ast n_{\mathrm{I.}}\ast {\mathrm{M.}}_{\mathrm{I.}}\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}$$

$${\mathrm{M.}}_{\mathrm{I.}}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{A.}}\ast v_{\mathrm{D.}urcHflus\mathrm{<figure-callout\; id="2"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}}{2\ast n_{\mathrm{I.}}\ast \pi \ast v_{\mathrm{A.}}\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}$$

$$\mathrm{S.}trmunGsGesHwindiGkeit:v_{\mathrm{S.}}=2\ast \pi \ast n_{\mathrm{I.}}\ast r$$

$${\mathrm{M.}}_{\mathrm{I.}}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{A.}}\ast v_{\mathrm{D.}urcHfluss}}{\frac{v_{\mathrm{S.}}}{r}\ast v_{\mathrm{A.}}\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}$$

$$\mathrm{F.}=\frac{{\mathrm{M.}}_{\mathrm{I.}}}{r}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{A.}}\ast v_{\mathrm{D.}urcHfluss}}{r\ast \frac{v_{\mathrm{S.}}}{r}\ast v_{\mathrm{A.}}\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{A.}}\ast v_{\mathrm{D.}urcHfluss}}{v_{\mathrm{S.}}\ast v_{\mathrm{A.}}\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}$$

(Quelle: https://www.easa.europa.eu/sites/default/files/dfu/EASA%20E110%20TCDS%20Issue%207%20LEAP-1A-1C.pdf; 25.10.2018, 10:25 Uhr). Damit ergibt sich für den Wirkungsgrad des Getriebes: η_Getriebe = (0,9825 * 0,9825² * 0,9825²)¹ = 0,9155. Für die Kraft F bedeutet dies: $$F=\frac{F_{Schub}\ast \mathrm{420,0}\frac{m}{s}\ast v_{Druchfluss}}{\mathrm{1625,0}\frac{m}{s}\ast \mathrm{420,0}\frac{m}{s}\ast \mathrm{0,90}*\mathrm{0,9155}}=\mathrm{0,00008233}\frac{s}{m}\ast F_{Schub}\ast v_{Durchfluss}$$

A special feature of STI engines is the gearbox to be used to reduce the impulse turbine speed to the speed of the low-pressure turbine. Since the speed of the low-pressure turbine is relatively high compared to that of a main rotor of a helicopter or a propeller of a turboprop aircraft, it can be assumed that a 1-stage gearbox is sufficient at this point. For example, the speed of the low pressure turbine in the CFM is LEAP-1A $3894\frac{1}{min}=64.9\frac{1}{s}$

(Source: https://www.easa.europa.eu/sites/default/files/dfu/EASA%20E110%20TCDS%20Issue%207%20LEAP-1A-1C.pdf; 25.10.2018, 10:25 am) . This results in the efficiency of the gearbox: η _gearbox = (0.9825 * 0.9825 ² * 0.9825 ² ) ¹ = 0.9155. For the force F this means: $$\mathrm{F.}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub}\ast 420.0\frac{m}{s}\ast v_{\mathrm{D.}rucHfluss}}{1625.0\frac{m}{s}\ast 420.0\frac{m}{s}\ast 0.90*0.9155}=0.00008233\frac{s}{m}\ast {\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHfluss}$$

$${\dot{m}}_{Kerosin}=F_{Schub}\ast v_{Durchfluss}\ast \mathrm{0,0000000010133}\frac{kg}{N\ast s}$$

The following can be determined for fuel consumption: $${\dot{m}}_{KerOsin}=\frac{\mathrm{F.}}{v_{\mathrm{S.}}}\ast \frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{0.00008233\frac{s}{m}\ast {\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHfluss}}{1625.0\frac{m}{s}}\ast 0.02$$

$${\dot{m}}_{KerOsin}={\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHfluss}\ast 0.0000000010133\frac{kG}{N\ast s}$$

Assuming the flow velocity v _s as a fixed variable, the fuel consumption can be determined depending on the required thrust.

als fix angenommen wird. $$P\left[PS\right]=P\ast 736\left[W\right]=P\ast 736\left[\frac{kg\ast m^2}{s^3}\right]=F_{Schub}\ast v_{Durchfluss}\ast 736\left[\frac{kg\ast m^2}{s^3}\right]$$

$$F_{Schub}\ast v_A*\frac{v_{Durchfluss}}{v_A}=2\ast \pi \ast n_I\ast M_I\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}$$

$$M_I=\frac{F_{Schub}\ast v_{Durchfluss}}{2\ast n_I\ast \pi \ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}$$

$$F=\frac{M_I}{r}=\frac{F_{Schub}\ast v_{Durchfluss}}{r\ast 2\ast n_I\ast \pi \ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}$$

$$F=\frac{F_{Schub}\ast v_{Durchfluss}}{2\ast \pi \ast r\ast 320\frac{1}{s}*\mathrm{0,90}*\mathrm{0,9155}}$$

$$F=\frac{F_{Schub}\ast v_{Druchfluss}}{2\ast \pi \ast r*320\frac{1}{s}\ast \mathrm{0,90}\ast \mathrm{0,9155}}=\frac{F_{Schub}\ast v_{Durchfluss}}{r}\ast \mathrm{0,00060363}\ast 1s$$

Since with jet engines it can be assumed that the radius of the impulse turbine is in most cases less than 0.779m due to the small dimensions of the engines, a fuel consumption calculation now follows, with the speed $n_{\mathrm{I.}}=320\frac{1}{s}$

is assumed to be fixed. $$\mathrm{P.}\left[\mathrm{P.}\mathrm{S.}\right]=\mathrm{P.}\ast 736\left[\mathrm{W.}\right]=\mathrm{P.}\ast 736\left[\frac{kG\ast m^2}{s^3}\right]={\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHfluss}\ast 736\left[\frac{kG\ast m^2}{s^3}\right]$$

$${\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{A.}}*\frac{v_{\mathrm{D.}urcHfluss}}{v_{\mathrm{A.}}}=2\ast \pi \ast n_{\mathrm{I.}}\ast {\mathrm{M.}}_{\mathrm{I.}}\ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}$$

$${\mathrm{M.}}_{\mathrm{I.}}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHflus\mathrm{<figure-callout\; id="2"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}}{2\ast n_{\mathrm{I.}}\ast \pi \ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}$$

$$\mathrm{F.}=\frac{{\mathrm{M.}}_{\mathrm{I.}}}{r}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHfluss}}{r\ast 2\ast n_{\mathrm{I.}}\ast \pi \ast {\mathrm{\eta }}_{Turbine}\ast {\mathrm{\eta }}_{Getriebe}}$$

$$\mathrm{F.}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHflus\mathrm{<figure-callout\; id="2"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}}{2\ast \pi \ast r\ast 320\frac{1}{s}*0.90*0.9155}$$

$$\mathrm{F.}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}rucHflus\mathrm{<figure-callout\; id="2"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}}{2\ast \pi \ast r*320\frac{1}{s}\ast 0.90\ast 0.9155}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHfluss}}{r}\ast 0.00060363\ast 1s$$

$${\dot{m}}_{Kerosin}=\frac{F_{Schub}\ast v_{Durchfluss}}{r^2}\ast \mathrm{0,00000000600442}\frac{kg*m^2}{N*s}$$

The following can then be determined for the fuel consumption: $${\dot{m}}_{KerOsin}=\frac{\mathrm{F.}}{v_{\mathrm{S.}}}\ast \frac{{\dot{m}}_{KerOsin}}{{\dot{m}}_{\mathrm{L.}uft}}=\frac{0.00060363\ast {\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHflus\mathrm{<figure-callout\; id="2"\; label="s"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">s</figure-callout>}}}{2\ast \pi \ast 320\frac{1}{s}\ast {\mathrm{<figure-callout\; id="2"\; label="r"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">r</figure-callout>}}^2}\ast 0.02$$

$${\dot{m}}_{KerOsin}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub}\ast v_{\mathrm{D.}urcHfluss}}{{\mathrm{<figure-callout\; id="2"\; label="r"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">r</figure-callout>}}^2}\ast 0.00000000600442\frac{kG*m^2}{N*s}$$

Assuming the speed n _I as a fixed variable, the fuel consumption can be determined as a function of the required thrust in relation to the radius r of the impulse turbine.

Schubkraft leistet und in der Impulsturbine einen Radius von r = 0,30m aufweist, etwa $\mathrm{2248,0}\frac{l}{h}$

Kerosin verbrauchen. Ein solches Triebwerk entspricht in etwa dem Leistungsspektrum eines konventionellen A320-Triebwerks, liegt im Verbrauch jedoch deutlich darüber.According to this general formula, an STI engine that operates at a flow rate of $v_{\mathrm{D.}urcHfluss}=288.0\frac{m}{s}26000N$

Makes thrust and in the impulse turbine has a radius of r = 0.30m, about $2248.0\frac{l}{H}$

Consume kerosene. Such an engine roughly corresponds to the performance spectrum of a conventional A320 engine, but its consumption is significantly higher.

einen Radius r = 0,48m besitzen, so würde es nur noch $\mathrm{878,0}\frac{l}{h}$

verbrauchen. Komplizierter wird dann jedoch die Gestaltung des Triebwerks, da sich mit zunehmendem Radius der Impulsturbine der Rauminhalt der Mantelkammer des Triebwerks verringert, solange die äußeren Abmessungen des Triebwerks sich nicht verändern. Abhängig vom jeweiligen Triebwerk ist die Länge des Radius' so auf ein bestimmtes Maß begrenzt.From this calculation it can be seen that an STI engine only leads to savings in fuel consumption with a radius length greater than r = 0.45m. Would an STI engine with 26000N thrust and $v_{\mathrm{D.}urcHfluss}=288.0\frac{m}{s}$

have a radius r = 0.48m, it would only $878.0\frac{l}{H}$

consume. The design of the engine then becomes more complicated, however, since the volume of the jacket chamber of the engine decreases as the radius of the impulse turbine increases, as long as the external dimensions of the engine do not change. Depending on the particular engine, the length of the radius is limited to a certain amount.

(Quelle: http://www.jet-engine.net/civtfspec.html; 25.10.2018, 23:20 Uhr) Nach Anwendung obiger Formel ergibt sich: $${\dot{m}}_{Kerosin}=\frac{F_{Schub,A330}\ast v_{Durchfluss}}{r^2}*\mathrm{0,00000000600442}\frac{kg*m^2}{N*s}$$

For wide-body aircraft such. B. an A330, however, can achieve relatively good savings. A Pratt & Wittney PW4168 engine, for example, has a fan diameter of approx. 2.5 m. (Source: http://www.jet-engine.net/civtfspec.html; 10/25/2018, 11:20 p.m.) It is therefore possible that the impulse turbine in the engine core accounts for 1.5 m of this. This would ultimately result in a radius r = 0.70 m if the casing of the impulse turbine is taken into account. A Rolls-Royce A330 engine, a so-called Trent 772-60, produces around 51172N when cruising. Its consumption is then included $3684\frac{l}{H}$

(Source: http://www.jet-engine.net/civtfspec.html; 25.10.2018, 11:20 p.m.) Applying the above formula results in: $${\dot{m}}_{KerOsin}=\frac{{\mathrm{F.}}_{\mathrm{S.}cHub,\mathrm{A.}330}\ast v_{\mathrm{D.}urcHfluss}}{{\mathrm{<figure-callout\; id="2"\; label="r"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">r</figure-callout>}}^2}*0.00000000600442\frac{kG*m^2}{N*s}$$

(Quelle: http://www.jet-engine.net/civtfspec.html; 25.10.2018, 23:38 Uhr) Daraus lässt sich die Durchflussgeschwindigkeit im Original-Triebwerk bestimmen: $$v_{Trent772-60}=\frac{\mathrm{316376,596}N}{\mathrm{897,191379}\frac{kg}{s}}=\mathrm{352,625}\frac{m}{s}$$

The flow velocity v flow can be derived from existing knowledge: A Trent 772-60B generates a maximum thrust of 316376,596N and an air _{flow rate of} $897.191379\frac{kG}{s}.$

(Source: http://www.jet-engine.net/civtfspec.html; October 25, 2018, 11:38 pm) From this, the flow rate in the original engine can be determined: $$v_{Trent772-60}=\frac{\mathrm{316376,596}N}{897.191379\frac{kG}{s}}=352.625\frac{m}{s}$$

$$A_{STI}={\left(\frac{\mathrm{2,53492}m}{2}\right)}^2\ast \pi -{\left(\frac{\mathrm{1,5}m}{2}\right)}^2\ast \pi =\mathrm{3,27968}{m}^2$$

In order to determine how much the flow rate would increase in this case in the STI engine, it is necessary to put the intersected areas of the turbine inlets in relation to one another. In the case of the Trent 772-60, this represents the entire area of the low-pressure turbine, while in the case of the STI engine, only the area that leads the air flow into the engine casing chamber may be calculated. For this purpose, the area of the turbine inlet minus the frontal area of the engine core is calculated for the STI engine. The engine core is set with a diameter of 1.5 m. It follows: $${\mathrm{A.}}_{Trent772-60}={\left(\frac{2.47396m}{2}\right)}^2\ast \pi =4.8070125{\mathrm{<figure-callout\; id="2"\; label="m"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">m</figure-callout>}}^2$$

$${\mathrm{A.}}_{\mathrm{S.}T\mathrm{I.}}={\left(\frac{2.53492m}{2}\right)}^2\ast \pi -{\left(\frac{1.5m}{2}\right)}^2\ast \pi =3.27968{\mathrm{<figure-callout\; id="2"\; label="m"\; filenames="DE102019130861B4\_0148.png"\; state="\{\{state\}\}">m</figure-callout>}}^2$$

The following applies to speed: $$v_{\mathrm{S.}}=v_{Trent772-60}\ast \frac{{\mathrm{A.}}_{Trent772-60}}{{\mathrm{A.}}_{\mathrm{S.}T\mathrm{I.}}}=516.85\frac{m}{s}$$

From this, the fuel consumption for the STI engine is determined: $${\dot{m}}_{KerOsin}=\frac{51172N*516.85\frac{m}{s}}{{\left(0.70m\right)}^2}\ast 0.00000000600442\frac{kG*m^2}{N*s}=0.3241\frac{kG}{s}$$

This corresponds to a consumption of: $${\dot{m}}_{KerOsin}=0.3241\frac{kG}{s}=0.3241\ast 3600\frac{kG}{H}=1166.76\frac{kG}{H}\approx 1458.45\frac{l}{H}$$

Compared to the conventional Trent 772-60, the consumption of an STI engine with the same power spectrum is approx. 40%. Similar savings potential can be assumed for aircraft of comparable size and comparable engines such as the Boeing 787, Boeing 777 or A350. For engines with relatively small dimensions, as z. For example, if they are used on an A340 or a Boeing 757, STI engines are poorly suited as replacements.

Figure list

• shows schematically an engine with a rotor as a drive unit.
• shows schematically an engine with a propeller as a drive unit.
• shows schematically an engine with a second jet engine as a drive unit.

Detailed description of the images:

shows schematically an engine with a rotor (rotor blades 13th ) as a drive unit, especially for a helicopter. The figure shows a top view in which elements that are shown in different levels can be seen together. There are two first jet engines 3 used, their mass flow 5 via an entrance area for the mass flow of the first jet engine 9 on the blades of the impulse turbine 1 be steered and put them in rotation. This rotation is due to the shaft or a matter of rotation 7th and from there on to the rotor, its rotor blades 13th preferably sit on the same shaft 7.

The mass flow of the fluid is via the exhaust port 11 diverted. The impulse turbine 1 lies in the plane of rotation of the impulse turbine 2 in front.

A fundamentally similar structure is used in shown, with the propeller blades here instead of the rotor 15th of the propeller are driven, which preferentially on with the blades of the impulse turbine on the same shaft 7th to sit.

shows a structure with a second jet engine as a drive unit, in which the mass flow 5 of the first jet engine via an entrance area for the mass flow 9 fed and on the impulse turbine 1 is directed. A rotation of the impulse turbine 1 is about the wave 7th on the fan of the second jet engine 17th is transmitted.

List of reference symbols

1

Impulse turbine

2

Plane of rotation of the impulse turbine

3

First jet engine

5

Mass flow

7th

Shaft / axis of rotation

9

Entrance area for the mass flow of the first jet engine

11

Exhaust opening

13th

Rotor blade

15th

Propeller blade

17th

Fan of the second jet engine

Claims (14)

Engine for an aircraft or rail vehicle, comprising: - at least one first jet engine (3) - at least one impulse turbine (1) comprising several blades mounted on a shaft (7), - at least one drive unit for generating propulsion of the aircraft or rail vehicle thereby characterized in that the first jet engine (3) generates a mass flow (5) of a fluid which is directed with a flow direction within a plane of rotation of the impulse turbine (1) on its blades, the shaft (7) of the impulse turbine (1) with the drive unit connected for a drive of the drive unit by the impulse turbine (1). Engine according to the preceding claim, characterized in that the blades of the impulse turbine (1) are configured for a transmission of a mass flow impulse of the fluid which has a direction essentially within the plane of rotation of the impulse turbine (2). Engine according to one or more of the preceding claims, characterized in that the blades have an essentially perpendicular surface to the plane of rotation of the impulse turbine (2). Engine according to one or more of the preceding claims, characterized in that a longitudinal axis of the first jet engine (3) is arranged in a plane parallel to the plane of rotation of the impulse turbine (2) or within the plane of rotation of the impulse turbine (2). Engine according to one or more of the preceding claims, characterized in that the shaft (7) of the impulse turbine (1) is connected directly to a shaft of the drive unit. Engine after one or more of the previous ones Claims 1 - 4th , characterized in that the shaft (7) of the impulse turbine (1) is connected to a shaft of the drive unit via a gear, in particular a spur gear and / or multi-stage gear. Engine according to one or more of the preceding claims, characterized in that the impulse turbine (1) is arranged in a turbine chamber which has at least one inlet area for the mass flow of the first jet engine (9) and at least one exhaust gas opening (11) and which preferably for a efficient guidance of the mass flow (5) on the blades of the impulse turbine (1) is configured. Engine according to one or more of the preceding claims, in particular for a helicopter, characterized in that the impulse turbine (1) drives a rotor and / or a propeller as a drive unit, the plane of rotation of the impulse turbine (2) being arranged parallel to a plane of rotation of the rotor and / or of the propeller and wherein the impulse turbine (1) and rotor and / or propeller have a common axis of rotation (7). Engine according to the preceding claim, characterized in that a longitudinal axis of the first jet engine (3) is arranged in a plane parallel to the plane of rotation of the impulse turbine (2) or within the plane of rotation of the impulse turbine (2). Engine after one or more of the previous ones Claims 1 - 3 , in particular comprising a turboprop engine, characterized in that the impulse turbine (1) drives a propeller as a drive unit, the plane of rotation of the impulse turbine (2) being arranged parallel to a plane of rotation of the propeller, the impulse turbine (1) and propeller having a common axis of rotation (7), wherein a longitudinal axis of the first jet engine (3) is preferably arranged in a plane perpendicular to the plane of rotation of the impulse turbine (2) in the direction of flight and the turbine chamber is configured for a deflection of the mass flow (5) so that it coincides with a flow direction is directed within the plane of rotation of the impulse turbine (1) on the blades. Engine after one or more of the previous ones Claims 1 - 3 comprising a jacketed engine fan, characterized in that the impulse turbine (1) drives the engine fan (17) as a drive unit, the plane of rotation of the impulse turbine (2) being arranged parallel to a plane of rotation of the engine fan (17), the impulse turbine (1) and engine fan ( 17) have a common axis of rotation (7), a longitudinal axis of the first jet engine (3) preferably being arranged in a plane perpendicular to the plane of rotation of the impulse turbine (2) in the direction of flight and the turbine chamber being configured for a deflection of the mass flow (5), see above that this is directed with a flow direction within the plane of rotation of the impulse turbine (2) on its blades. Engine after one of the previous ones Claims 10 or 11 , characterized in that the turbine chamber comprises an arcuate deflection of the mass flow. Engine after Claim 11 , characterized in that the turbine chamber is arranged within the casing of the engine fan and is configured for a hermetic separation of the mass flow (5) from a fluid flow within the engine fan. Engine after one of the previous ones Claims 11 or 13th , characterized in that there is an intake port for air of the first jet engine (3) arranged on a hub of the engine fan.

Images