engines

Abstract

There is provided an engine, comprising a core air inlet configured to receive air and generally centred on a longitudinal axis of the engine, a bypass air inlet configured to receive air, a first heat exchanger arrangement arranged downstream of the core inlet and configured to cool air received in the core inlet, one or more turbomachinery components configured to receive air and arranged downstream of the first heat exchanger arrangement, and centred on the longitudinal axis, at least one rotor centred on the longitudinal axis and having a radial direction perpendicular to the longitudinal axis, a first flow path arranged to extend through the core inlet, the first heat exchanger arrangement, the one or more turbomachinery components and an inner radial portion of the at least one rotor, and a second flow path arranged to extend through the bypass inlet and an outer radial portion of the at least one rotor, the second flow path arranged to bypass the first heat exchanger arrangement, and the one or more turbomachinery components.

Description

[0001] Engines

[0002] Technical field

[0003] The present disclosure relates to engines, in particular precooled engines, such as of the type which may be used in aerospace applications, and to methods of operating such engines. The present disclosure also relates to aircraft, flying machines and aerospace vehicles comprising such engines.

[0004] Background

[0005] Efficient air-breathing propulsion for flight from static (i.e. take off) to high Mach numbers (i.e. Mach 2.5 - Mach 5 or above) requires the integration of several engine types that each operate in separate flight speed regimes. For example, an air-breathing turbomachinery-based engine, such as a turbojet or a turbofan (collectively known as gas turbine jet engines), is required to propel an aircraft from static at ground level to low Mach numbers (up to about Mach 2). The aircraft must then transition to an alternative engine such as a ramjet or rocket engine, that is capable of providing thrust at higher Mach numbers. The integrated engine system requires carefully managed transitions between the different modes of operation at the different respective Mach numbers, as thrust falls in one mode while thrust in another mode rises.

[0006] The Mach number at which these transitions between different modes of operation occurs is largely a result of material limitations (in particular material temperature limitations) and the decreasing ability of a gas turbine to generate thrust at higher flight Mach numbers. Gas turbine jet engines have very high temperatures at the exit of the high-pressure compressor and at the exit of the combustor (turbine inlet), even at low Mach numbers. As the flight speed increases (e.g. Mach 2 and above), the gas turbine engine becomes temperature limited because of material limitations and poor thrust performance due to the increased stagnation temperature of the freestream air entering the engine. In particular, the increase in stagnation temperature results in the temperature ratio between the engine inlet and the high-pressure compressor (HPC) outlet reducing, which directly and proportionally reduces the gas turbine engine's thrust output.

[0007] These limitations in conventional gas turbine engines mean that the transition to ramjet operation is required at a flight speed that is not high enough for efficient ramjet performance (i.e. a flight speed for which the ram pressure is too low for producing good ramjet thrust output). Consequently, the transition between the modes of operation occurs at a flight speed where neither mode is optimised, thereby resulting in a gap / deficiency in overall thrust performance in which the propulsion system as a whole struggles to provide sufficient thrust to continue acceleration through the transition phase from gas turbine to ramjet.

[0008] One known technique for addressing this reduction in thrust at the transition speed is to cool the freestream air that is ingested into the engine (known as precooling). This allows the gas turbine to continue providing acceptable thrust output at flight speeds that are beyond its typical operating envelope, thereby delaying the transition point so that the transition to ramjet operation occurs at a higher Mach number (at which the ram pressure will naturally be higher, thereby allowing the ramjet to produce a better thrust output). In other words, pre-cooling the freestream air entering the engines allows the various engine modes to be better optimised, such that the need for the ramjet is only required for later stages of acceleration to a desired cruise speed.

[0009] Therefore, to achieve good thrust performance across a wide range of Mach numbers (e.g. up to Mach 5), known solutions involve the use of a precooled turbofan architecture with reheat. An advantage of such engines is that at low speed, the bypass stream is pressurised using the low- pressure fan which provides for good fuel consumption at subsonic speeds. At the same time, such engines also allow for higher maximum achievable flight Mach number under turbofan operation before transitioning to ramjet, due to the benefits of pre-cooling the inlet air which allows the gas turbine to operate at higher than usual flight speeds.

[0010] An issue with known pre-cooled turbofan architectures as discussed above relates to gas turbine engine choking at flight Mach numbers around Mach 2.5. Traditional turbofans (without pre-coolers) are designed to operate only over limited Mach ranges (e.g. from zero up to Mach 2.2). At higher Mach numbers, the mixer that matches the bypass and core flows to maximise thrust can choke, since the mixer is designed for matching requirements at lower speeds (e.g. only up to Mach 2.2) and is therefore not optimised for mixing at higher speeds. At the same time, higher flight Mach numbers (e.g. Mach 2.5 and above) lead to increased flow through the bypass, which can lead to choking at the exit of the fan guide vanes and bypass pressure loss due to blockage caused by the fan rotors and stators. The pressure loss in the bypass at such high flight Mach numbers results in insufficient pressure for efficient ramjet operation. Choking in the bypass compressor fan blades and guide vanes also limits the mass flow rate through the bypass, since the high mass flow rate resulting from the higher flight Mach numbers reaches a limit due to the flow through the compressor reaching a velocity of Mach 1, at which point the flow is said to be choked because it cannot be increased further. One known solution for mitigating pressure losses in the bypass of a gas turbine engine is to increase the bypass ratio, thereby allowing more airflow through the bypass. However, a drawback of this solution is that this invariably requires a larger engine, thereby increasing weight and reducing the specific fuel efficiency. Moreover, an airframe designed to support such a larger engine is required to be stronger, thereby increasing the overall weight of the airframe. In addition, a larger bypass ratio results in a lower bypass pressure, which deteriorates the efficiency of the after burner.

[0011] There is therefore a need for an improved pre-cooled combined turbofan and ramjet architecture that facilitates a smooth transition from turbofan operation at lower flight speeds to ramjet operation at higher flight speeds.

[0012] The present disclosure seeks to address this and other problems encountered in the prior art.

[0013] Summary

[0014] According to a first aspect of the present disclosure, there is provided an engine comprising a core air inlet configured to receive air and generally centred on a longitudinal axis of the engine, a bypass air inlet configured to receive air, a first heat exchanger arrangement arranged downstream of the core inlet and configured to cool air received in the core inlet, and one or more turbomachinery components configured to receive air and arranged downstream of the first heat exchanger arrangement and centred on the longitudinal axis. The engine comprises at least one rotor centred on the longitudinal axis and having a radial direction perpendicular to the longitudinal axis, a first flow path arranged to extend through the core inlet, the first heat exchanger arrangement, the one or more turbomachinery components and an inner radial portion of the at least one rotor, and a second flow path arranged to extend through the bypass inlet and an outer radial portion of the at least one rotor. The second flow path is arranged to bypass the first heat exchanger arrangement, and the one or more turbomachinery components.

[0015] The at least one rotor may be arranged downstream of the one or more turbomachinery components. Alternatively, the at least one rotor may be arranged upstream of the one or more turbomachinery components and downstream of the first heat exchanger.

[0016] In some embodiments, the core inlet is defined by a generally cylindrical inner casing within which the first heat exchanger arrangement is arranged, and the bypass inlet defined by a generally cylindrical outer casing arranged concentric with and radially outside the inner casing. In such embodiments, the first flow path extends through the inner casing and the second flow path extends between the inner and outer casings.

[0017] In some embodiments, the axial extent of the inner casing in a direction away from the turbomachinery components is greater than an axial extent of the first heat exchanger arrangement in the direction away from the turbomachinery components.

[0018] In some embodiments, the engine further comprises a main inlet upstream of the core inlet and bypass inlet, wherein the first and second flow paths extend from a common flow path that extends through the main inlet.

[0019] In some embodiments, the at least one rotor comprises a plurality of blades angularly spaced apart from each other and each arranged to extend generally along the radial direction of the rotor, wherein each of the blades has a root corresponding to the inner radial portion of the rotor and a tip corresponding to the outer radial portion of the rotor, the root being arranged closer to the longitudinal axis than the tip such that the first flow path is arranged to extend adjacent the roots of the blades and the second flow path is arranged to extend adjacent the tips of the blades. The root and tip of each blade may be made of the same material.

[0020] In some embodiments, the one or more turbomachinery components comprise at least one compressor downstream of the first heat exchanger arrangement and at least one combustion chamber.

[0021] In some embodiments, the one or more turbomachinery components comprise at least one combustion chamber and at least one turbine downstream of the at least one combustion chamber, wherein the at least one turbine is coupled to a respective one of the at least one rotors via a shaft. The at least one combustion chamber may comprise an air-breathing combustion chamber configured to receive compressed air for combustion, wherein the combustion chamber is configured to mix the received compressed air with fuel, and combust the compressed air and fuel mixture to produce exhaust gas.

[0022] In some embodiments, the at least one rotor is arranged downstream of the turbine such that the inner radial portion of the at least one rotor is arranged to receive the exhaust gas to thereby cause rotation of the rotor about the longitudinal axis as a result of the exhaust gas flowing through the inner radial portion of the rotor, and rotation of the rotor caused by the exhaust gas flowing through the inner radial portion of the rotor causes the outer portion of rotor to compress air received by the bypass inlet flowing along the second flow path.

[0023] In some embodiments, the at least one turbine is arranged to rotate as a result of the exhaust gas flowing through the at least one turbine, and wherein rotation of the at least one turbine in turn causes rotation of the respective at least one rotor via the respective shaft, and rotation of the rotor caused by rotation of the turbine causes the inner radial portion of the rotor to compress air receiving in the core inlet along the first flow path, and causes the outer portion of rotor to compress air received by the bypass inlet flowing along the second flow path.

[0024] In some embodiments, the at least one rotor comprises two rotors each arranged along and centred on the longitudinal axis of the engine and arranged to rotate independently about the longitudinal axis. Optionally, the two rotors are arranged to rotate in opposite directions. The two rotors may be adjacent along the longitudinal axis. In some embodiments, the engine does not comprise stator vanes positioned in between the two rotors along the longitudinal direction.

[0025] In some embodiments, the engine comprises an exhaust nozzle, wherein the first and second flow paths are arranged to flow from their respective air inlets to the exhaust nozzle. The engine optionally further comprises a reheat system arranged downstream of the at least one rotor, wherein the first and second flow paths both extend through the reheat system.

[0026] According to a second aspect of the present disclosure, there is provided a method of operating an engine according to the first aspect. The method comprises causing a first portion of air to flow along the first flow path such that at least a portion of the air received by the core inlet is cooled in the first heat exchanger arrangement, causing, using the one or more turbomachinery components, combustion of the first portion of air to produce a first exhaust gas, causing at least a portion of the first exhaust gas to flow through an inner radial portion of the at least one rotor, thereby driving rotation of the rotor, causing the first exhaust gas to exit the engine through an exhaust nozzle to produce a first amount of thrust, causing a second portion of air, different from the first portion of air, to flow along the second flow path such that at least a portion of the air received by the bypass inlet is configured to extend through the outer portion of the one or more rotors, compressing, using the outer portion of the of the one or more rotors, at least a portion of the air received by the bypass inlet, combusting, using a reheat system, the compressed air to produce a second exhaust gas, and causing the second exhaust gas to exit the engine through the exhaust nozzle to produce a second amount of thrust.

[0027] In some embodiments, when a flight speed of the engine reaches a first threshold speed, the method comprises reducing an amount of fuel supplied to the one or more turbomachinery components to reduce the first amount of thrust produced by the engine. Optionally, when a flight speed of the engine reaches a second threshold speed that is greater than the first threshold speed, the method comprises stopping fuel flow to the one or more turbomachinery components such that the first amount of thrust becomes zero. Optionally, when the flight speed of the engine is greater than or equal to the second threshold speed, the first portion of air flowing along the first flow path drives rotation of the one or more turbomachinery components, and wherein the method further comprises generating electrical power using the rotation of the one or more turbomachinery components.

[0028] Optionally, when the flight speed of the engine is greater than or equal to the second threshold speed, the method comprises preventing rotation of the one or more turbomachinery components to reduce the amount of the first portion of air flowing along the first flow path.

[0029] In some embodiments, when the flight speed of the engine is greater than or equal to the second threshold speed, the second portion of air flowing along the second flow path drives rotation of the one or more rotors, and the method further comprises generating electrical power using the rotation of the one or more rotors.

[0030] According to a third aspect of the present disclosure, there is provided an aircraft, flying machine, or aerospace vehicle comprising an engine according to the first aspect.

[0031] Brief description of the drawings

[0032] The present disclosure may be carried out in various ways and embodiments of the disclosure will now be described by way of example with reference to the accompanying drawings, in which:

[0033] Figure 1 shows a cross-sectional schematic view of an engine according to the prior art;

[0034] Figure 2a shows a cross-section schematic view of an engine according to the present disclosure; Figure 2b shows a cross-section schematic view of an engine according to the present disclosure; Figure 3 shows a cross-section schematic view of an engine according to the present disclosure; and Figure 4 shows a rotor blade according to the present disclosure.

[0035] Detailed description of the drawings

[0036] Figure 1 (prior art) shows a cross-sectional schematic view of a precooled turbomachinery-based (i.e. gas turbine) engine 100. The engine has an air inlet 110, a heat exchanger 120, a low-pressure compressor / fan 130, a high-pressure compressor 140, a combustion chamber 150, a high-pressure turbine 160 and a low-pressure turbine 170. The components of the engine are arranged generally along a longitudinal axis indicated in the figure by the dashed line. The components are arranged generally inside an outer casing 102 of the engine. The outer casing is generally cylindrical and is centred on the longitudinal axis of the engine. The high-pressure compressor 140, the combustion chamber 150, the high pressure turbine 160 and the low pressure turbine may collectively be referred to in reference to this figure as turbomachinery components. As shown in the figure, between the turbomachinery components and the outer casing 102, there is arranged a generally cylindrical inner casing 104. The annular gap between the outer casing 102 and the inner casing 104 provides the bypass flow path for the engine. In other words, at least a portion of air entering the engine 100 through the inlet 110 will bypass the turbomachinery components (in the engine core and situated inside the inner casing 104) will flow through the gap between the inner and outer casings. The remainder of the air will flow through the engine core (the turbomachinery components) inside the inner casing 104.

[0037] The low-pressure compressor / fan has an outer radial portion 130a and an inner radial portion 130b. The inner radial portion is closer to the longitudinal axis relative to the outer radial portion. As would be appreciated by the skilled person, a low-pressure compressor / fan as is common in turbofan engines comprises a plurality of blades angularly spaced apart from one another and each arranged to extend from the centre of the compressor / fan at the longitudinal axis outwards in a radial direction perpendicular to the longitudinal axis. Each blade has a tip at the outer radial portion 130a and a root at the inner radial portion 130b. As shown in figure 1, the engine 100 has a first flow path 106 that extends from the air inlet 110 to one or more exhaust nozzles (not shown in the figure) downstream of the turbomachinery components via the heat exchanger 120, the inner radial portion of the low pressure compressor 130b and the turbomachinery components in the core of the engine. The engine also has a second flow path 108 that extends from the air inlet to one or more exhaust nozzles via the heat exchanger 120 and outer radial portion of the low pressure compressor / fan 130a, and bypassing the turbomachinery components in the core of the engine. The first and second flow paths 106 and 108 are arranged to extend generally parallel to the longitudinal axis, with the first flow path passing through the inside of the inner casing 104 and the second flow path passing in the gap between the inner casing 104 and the outer casing 102.

[0038] When in use, air passing into the engine 100 according to the prior art will first flow through a heat exchanger 120, which is configured to cool the air received in the inlet 110 before said air is compressed by the low-pressure compressor / fan 130. In more detail, a portion of the air received by the inlet 110 will flow along the first flow path, which will be cooled in the heat exchanger 120 before passing through the inner radial portion 130b of the low pressure compressor / fan before flowing into the turbomachinery components in the engine core. Another portion of the air received by the inlet 110 will flow along a second flow path, which will be cooled in the heat exchanger 120 before being compressed by the outer radial portion 130a of the low pressure compressor and bypassing the engine core by flowing through the gap between the inner casing 104 and the outer casing 102.

[0039] As is known in the art, cooling air before the air is received in the engine core provides specific advantages to the thrust output of the engine's core at high flight speeds. This is because the stagnation temperature of the air at the low pressure compressor and other turbomachinery components is reduced, thereby allowing the engine to operate at higher speeds before material temperature limitations of the engine's components are reached.

[0040] However, an issue with known engine arrangements according to figure 1 is that at least some of the air flowing along the second flow path (i.e. the bypass air) is also cooled. Considering the bypass ratio of an engine, which is the ratio between the mass flow rate of the bypass stream to the mass flow rate entering the engine core, which may typically between 0.5 and 1, around 30-50% of the heat transferred by the heat exchanger would be in air that flows through the bypass. In some scenarios, it may be advantageous to cool the bypass air, since this may lead to a higher pressure ratio across the fan which may improve thrust output. However, in many situations, cooling the bypass flow has an insignificant and often negligible thrust benefit. In fact, at certain conditions where the use of an after burner (i.e. a reheat system 190) is required, the lower temperature of the cooled bypass air would require more fuel flow in the reheat system to achieve a required after burner flame temperature, thus reducing the fuel efficiency. Cooling of the bypass air is therefore a wasted use of the heat-sink capacity of the heat exchanger 120, since, in turbofan operation, the thrust benefits of cooling the bypass air before the air reaches the low pressure compressor / fan 130 are negligible, and, in ramjet operation (in which the bypass air is burnt using a reheat system 190 in the exhaust of the engine), the bypass air is cooled only to be reheated again when it is burnt by the reheat system 190. Thus, either in turbofan operation (at low flight speeds), or in ramjet operation (at higher flight speeds), engine arrangements according to Figure 1 have less-than-optimal efficiency with regards to the heat sink capacity of the heat exchanger, as well as a potential for reduced fuel efficiency.

[0041] Turning to Figures 2a and 2b, cross-section schematic views of an engine 200 according to the present disclosure are depicted. Figures 2a and 2b show alternative arrangements of the core and bypass inlets, as discussed below. The engine comprises a core inlet 210, a bypass inlet 212, a heat exchanger 220, a compressor 240 (which may comprise multiple compressor stages, e.g. at least one low pressure stage and at least one high pressure stage), a combustion chamber 250, a turbine 260 (which may comprise multiple turbine stages, e.g. at least one high pressure turbine and at least one low pressure turbine) and one or more rotors 270 configured to rotate about the longitudinal axis. The core inlet 210 and bypass inlet 212 may be downstream of a larger main engine inlet (as can be seen for example in figure 2b) which is defined by an opening to a nacelle that contains all the components of the engine 200. The main engine inlet is the inlet that receives all airflow into the engine, which is then "split" downstream of the main engine inlet into the core flow and bypass flow. In other words, the main airflow received in the main engine inlet is split into the core and bypass flow at the point that the flow reaches the core inlet 210 and bypass inlet 212.

[0042] The arrangement and the operation of the rotors is described in more detail below. As with the engine shown in figure 1, the components of the fan are arranged generally along a longitudinal axis indicated in Figure 2a or 2b by the dashed line. The components are arranged generally inside an outer casing 202 of the engine, which is generally cylindrical and centred along the longitudinal axis of the engine. In the arrangement shown in Figure 2a, the core inlet 210 is configured to receive freestream air and is arranged along and generally centred on the longitudinal axis. The bypass inlet 212 is also configured to receive freestream air and is also arranged along and generally centred on the longitudinal axis. The bypass inlet 212 is concentric with the core inlet 210 and is positioned radially outside of the core inlet, where the radial direction is any direction perpendicular to the longitudinal axis. In other words, the bypass inlet has a larger radius than the core inlet. In some examples, the core inlet 210 has a generally annular shape with an inner radius and shape defined by a cone 230, and an outer radius and shape defined by an inner casing 204. In some examples, the bypass inlet 212 has a generally annular shape with an inner radius defined by the inner casing 204 and the outer radius defined by the outer casing 202. The outer and inner casings 202 and 204 may have generally cylindrical shapes that are centred along the longitudinal axis. The annular gap between the outer casing 202 and the inner casing 204 provides the bypass flow path for the engine 200. In other words, at least a portion of the freestream air entering the engine at the main engine inlet will then flow through by bypass inlet 212 and will bypass the components situated inside the inner casing 204. This air is referred to as the "bypass flow" or "bypass air" as it flows through the bypass duct of the engine and bypasses the engine's core. As described in more detail below, the bypass flow is pressurised by an outer radial portion of the one or more rotors 270 that is driven by the core flow, wherein the rotor(s) 270 acts as a fan that pressurise the bypass flow to produce thrust. Another, different portion of the freestream air entering the engine at the main engine inlet will then flow through core inlet 210 and will flow inside the inner casing 204 and through the engine core, which comprises the heat exchanger 220, compressor 240, combustion chamber 250, turbine 260 and inner radial portion of the rotor 270. This air is referred to as the "core flow" or "core air" as it flows through the engine's core, where it is compressed by the compressor 240 before being mixed with fuel and burnt in the combustion chamber 250. As shown in the figure, the core flow is directed through the heat exchanger 220 in an inwardly radial direction. The inner casing may be designed to cause the air received in the core inlet to flow radially through the heat exchanger before flowing generally axially through the rest of the core turbomachinery components. Generally speaking, in the example depicted in Figure 2a, the inner casing 204 acts as a splitter which splits the air received in the main engine inlet into the core flow and the bypass flow. In other words, the inner casing acts as the boundary that separates the core and bypass inlets and thus separates the core and bypass flows.

[0043] Figure 2b illustrates an engine 200 with the same components as shown in Figure 2a which has an alternative arrangement of the core and bypass inlets. In this example, air is received into the engine via a main engine inlet 201. The flow is air is then split downstream of the main engine inlet 201 into a first flow path 206 (the core flow) and a second flow path 208 (the bypass flow). The splitting of the air flow into the first and second flow paths can be seen in figure 2b as indicated by the dashed arrows. The first flow path 206 flows through a core inlet 210 and radially through heat exchanger 220 in the same manner as described above in relation to Figure 2a. The second flow path 208 flows through a bypass inlet 208 and into the bypass duct in the same manner as described above. Thus, the engine shown in Figure 2b also has first and second flow paths in the same manner as the engine shown in Figure 2a, where the first flow path extends through a core inlet (configured to receive air that will flow into the core), the heat exchanger and one or more turbomachinery components and an inner radial portion of the rotors 270, and where the second flow path extends through a bypass inlet 212 and an outer radial portion of the rotors 270, where the second flow path bypasses the heat exchanger 220 and the turbomachinery components in the core of the engine. In each of Figures 2a and 2b, the first and second flow paths may start as a common flow path of air that is received into the main inlet of the engine. This common flow path then be split into the core and bypass flow paths as shown in Figure 2b, or may be split according to a boundary (the inner casing) as shown in Figure 2a.

[0044] In some examples, the engine 200 may be designed to have a bypass ratio of between 0.3 and 2, or between 10 and 15, or anywhere in between.

[0045] With reference to Figure 2a and 2b, compressor 240, combustion chamber 250, turbine 260 may be collectively referred to as turbomachinery components. Some embodiments of the present disclosure may involve a single-stage compressor or multiple-stage compressor (at least one low pressure compressor and at least one high pressure compressor). Generally, the turbine 260 will have a number of stages that correspond to the compressor stages. For example, if the compressor is a single-stage compressor, the turbine may be a single-stage turbine that is connected to the compressor via a shaft so that the turbine can drive the shaft. In other examples, the compressor may comprise a low pressure stage and a high pressure stage and the turbine correspondingly comprises a low pressure stage (connected via a low pressure shaft to the low pressure compressor) and a high pressure stage (connected via a high pressure shaft to the high pressure compressor).

[0046] As can be seen in Figure 2a and 2b, the heat exchanger 220 is arranged downstream of the core inlet 210 (downstream of the airflow through the engine). In the example depicted in Figure 2a, the heat exchanger 220 is positioned inside of the inner casing 204 and is therefore physically separated from freestream air that is received in the bypass inlet 212 (i.e. the bypass flow). This means that the heat exchanger 220 is arranged such that is configured to cool only the air received in the core inlet, i.e. the core air flow. The arrangement of the heat exchanger 220 is such that no air that is received in the bypass inlet (i.e. none of the bypass flow) is cooled by the heat exchanger. In the example depicted in Figure 2b, the heat exchanger is also positioned downstream of the core inlet such that air cooled by the heat exchanger 220 flows only into the core of the engine and not into the bypass flow. In other words, air that is cooled in the heat exchanger 220 is then physically separated from the bypass flow by at least the inner casing 204 in either of the examples shown in Figure 2a or 2b. In either example, none of the air received in the bypass inlet 212 is cooled in the heat exchanger. Therefore, the engine 200 according to the present disclosure is particularly advantageous over some engines according to the prior art as shown in Figure 1, since the cooling capacity of the heat exchanger is directed to the core air flow only. This avoids bypass air being unnecessarily cooled (which, as noted above, achieves little to no benefit), meaning the capability of the heat exchanger to cool air passing through it is focused on the core air flow only. In this way, the cooling capacity of the heat exchanger is not unnecessarily wasted by cooling bypass air which does not need to be cooled. The arrangement of the heat exchanger being downstream of the core inlet in this way therefore provides improved core air cooling efficiency for the engine. Improving the cooling efficiency of the engine by only cooling that air which needs to be cooled (the cool air) can enable cooling budget to be freed up for other aspects of an aircraft carrying the engine.

[0047] The heat exchanger 220 may be any conventional heat exchanger that is suitable for cooling air received in an engine inlet as known in the art. The heat exchanger may transfer heat from the freestream air to any suitable heat sink. In some examples, as would be known by the skilled person, the heat exchanger may transfer heat from the freestream air to a working fluid, which may be fuel used in the engine itself. In these examples, improving the efficiency of the precooling of the freestream air in the manner disclosed, it is possible to reduce the heat load into the engine's fuel (since only the air received in the core needs to be cooled, rather than all of the air received in the engine). This means that some of the heat capacity of the fuel may be used for cooling elsewhere in the engine, for example cooling the reheat system, or in the aircraft more generally (for example cooling the electronic systems on the aircraft or cooling the airframe of the aircraft, which becomes increasingly important at increasingly higher flight speeds). A further advantage achieved by the engine shown in Figure 2a or 2b is that a smaller heat exchanger can be used compared to heat exchangers used in engines according to the prior art. This is because the heat exchanger is required to cool less air (only the core air flow), and the heat exchanger is positioned within the core inlet 210. Using a smaller heat exchanger reduces the overall size and weight of the engine, which leads to various benefits including fuel increased fuel efficiency.

[0048] After being cooled by the heat exchanger, the core air flow received in the core inlet 210 then flows through one or more compressor stages (the compressor 240) which compress the pre-cooled core flow, before the compressed core flow enters the combustion chamber 250. The operation of a gas turbine compressor, including the operation of multiple compressor stages, is well known to the skilled person. Similarly, the operation of a gas turbine combustion chamber is also well known to the skilled person. In brief, the core flow that has been cooled by the heat exchanger 220 and compressed by the compressor 240 is then mixed with fuel (optionally fuel that has been heated in the heat exchanger due to a heat exchange between the freestream core air flow and the fuel) in the combustion chamber and burnt to produce high pressure exhaust gases. The high-pressure exhaust gases exit the combustion chamber and flow through the turbine 260. The flow of the exhaust gases through the turbine 260 drives rotation of the turbine, as is well known by the skilled person. The turbine 260 is connected via one or more shafts to the one or more compressor stages, such that rotation of the turbine (which is driven by the flow of exhaust gases through the turbine) drives the compressor stages 240.

[0049] Downstream of the one or more turbomachinery components (i.e. the compressor stages 240, the combustion chamber 250, and the turbine 260) are one or more rotors 270, the structure and operation of which will now be described. Figures 2a and 2b depict two rotors 270, but embodiments of the invention may comprise a single rotor, or more than two rotors. In the following description, a single rotor will be described. Embodiments which comprise two rotors are described in more detail below.

[0050] A rotor 270 is centred on the longitudinal axis and is configured to rotate about the longitudinal axis. The rotor has a radial direction that is perpendicular to the longitudinal axis. The rotor has an outer radial portion 270a and an inner radial portion 270b, where the inner radial portion is closer to the longitudinal axis relative to the outer radial portion. As shown in Figure 2a or 2b, the outer radial portion of the rotor 270a is arranged radially between the inner and outer casings, such that the outer radial portion 270a corresponds to the bypass flow. In contrast, the inner radial portion 270b is arranged radially within the inner casing 204, such that the inner radial portion 270b corresponds to the core flow.

[0051] As with the low pressure compressor / fan depicted in Figure 1 and described above, the rotor 270 comprises a plurality of blades that are angularly spaced apart from one another and each arranged to extend from the centre of the rotor (at the longitudinal axis) outwards in a generally radial direction perpendicular to the longitudinal axis. An example rotor blade is depicted in Figure 4. The blade 400 comprises an outer radial portion 410 (i.e. a tip) and an inner radial portion 420 (i.e. a root). The tip and the root may be separated by a separating portion 415 which delineates the inner radial portion 420 from the outer radial portion 410. The separating portions of each blade may collectively provide a seal between the core flow path and the bypass flow since, in operation, there may be a pressure differential between the core and bypass flows at the point of the core turbine / bypass fan. For example, the separating portions of each blade may provide a leaf seal that separates the core flow from the bypass flow and that prevents leakage of the core flow into the bypass flow and vice versa. In other embodiments, the separating portions of each blade may be structurally arranged to provide any other suitable form of seal that is apparent to the skilled person and that separates the core and bypass flows when they pass through the rotor 270.

[0052] The blade may further comprise a connecting portion 430 that connects the blade 400 to a rotating hub or shaft (not shown) positioned along the longitudinal axis. The connecting portion 430 connects the blade 400 to the hub via the inner radial portion 420. The outer radial portion 410 of each blade in general corresponds to the outer radial portion 270a of the rotor 270 as a whole. Similarly, the inner radial portion 420 of each blade in general corresponds to the inner radial portion 270b of the rotor 270 as a whole.

[0053] Referring back to Figure 2a or 2b, engine 200 has a first flow path 206 that extends from the core inlet 210 to one or more exhaust nozzles (not shown in the Figure) via the heat exchanger 220, the turbomachinery components in the core of the engine, and via the inner radial portion (270b) of the rotor 270. Along the first flow path 206, the inner radial portion 270b is arranged so as to receive exhaust gas from the turbomachinery components, and the exhaust gas flows through the inner radial portion 270b which causes rotation of the rotor about the longitudinal axis as a result of the exhaust gas interacting with the inner radial portions 420 of the rotor blades 400. Thus, when the engine is operating as a turbofan, the inner radial portion acts as a turbine which is driven by the flow of exhaust gases through the inner radial portion 207b of the rotor 270.

[0054] Referring also to Figure 2a or 2b, the engine 200 also has a second flow path 208 that extends from the bypass inlet 212 to one or more exhaust nozzles via the outer radial portion 270a of the rotor. The second flow path bypasses the heat exchanger 220 and the turbomachinery components in the core of the engine. Along the second flow path 208, the outer radial portion 270a is arranged so as to receive the freestream bypass air flow that is received in the bypass inlet 212. As a result of the exhaust gases driving rotation of the inner radial portion 270b of the rotor, the outer radial portion 270a is driven to rotate since the outer radial portion and the inner radial portion of each rotor blade is connected (and in some examples the inner and outer portions are integral, meaning they formed of the same component). Thus, in operation of the engine 200 at low flight Mach numbers, rotation of the outer radial portion 270a (and more generally, of the rotor 270 as a whole) causes the freestream bypass air flowing into the outer radial portion 270a to be compressed, thereby producing thrust in the manner that a conventional turbofan produces thrust. Thus, the outer radial portion of the rotor acts as a fan that compresses the bypass flow, whilst the inner radial portion acts as a turbine that is driven by the exhaust gases produced in the engine's core. Referring back to Figure 4, the outer radial portion 410 may be considered to be a fan blade, since it is arranged to pressurise bypass air, whilst the inner radial portion 420 may be considered to be a turbine blade, since it is arranged to be driven by the core exhaust gas flow.

[0055] As mentioned above, each portion of the blades 400 of the rotor may be integral to each other. In other words, the portions of the blade 400 are collectively made of a single component, rather than separate components that are joined together. For example, each blade may be made of a singlecrystal super alloy that includes nickel, rare earth materials, and / or any other suitable material that would be known to the skilled person. In some examples, the blades may be made of a ceramic matrix composite. Any suitable manufacturing technique known to the skilled person may be used to manufacture the blades, such as investment casting, or the blades may be grown as a single crystal. Importantly, each blade is made of a material and is made using a manufacturing technique that is suitable for withstanding the temperatures typically seen in by a conventional turbine in a gas turbine engine.

[0056] Forming each blade of a single component is advantageous because it increases the rigidity and strength of the blade, which will be under a very high amount of stress in operation (due to high rotation speeds). In addition, at least the inner radial portion 420 of each blade is required to withstand very high temperatures (in some examples up to 1300K) due to the very hot exhaust gases flowing through the core (inner radial portion) of the rotor. As such, manufacturing blades for the rotor(s) of a single component for the purpose of improved strength and rigidity has the added advantage that the outer radial portions of the blades are also necessarily capable of withstanding very high (in some examples up to 1300K) temperatures. A stagnation temperature of 1300K at the outer radial portion of the blades corresponds to a flight speed of Mach 5, meaning that, because the rotor blades are made of a single piece of material and are required to withstand exhaust gas temperatures of 1300K, the rotor blades are also capable of withstanding the freestream bypass stagnation temperatures that are seen at Mach 5. This is another particular advantage achieved by some embodiments in which the bypass fan and core turbine blades are integral (due to the bypass fan, i.e. the outer radial portion of the rotor 270, being positioned at the rear of the engine and as part of the turbine blades, rather than being positioned at the front of the engine integral with the compressor blades as is the case in conventional turbofan designs). This advantageously means that there is no requirement for the bypass freestream air to be cooled before it reaches the rotor blades, even at hypersonic speeds where the stagnation temperature exceeds the material limits of conventional fan blade material. In some embodiments, the engine may be configured to reach a flight Mach number of Mach 6. At such a flight speed, the stagnation temperature is approximately 1620K. Thus, in some embodiments, the rotor blades are made of suitable materials that have a temperature limit greater than or equal to 1620K such that the inner portions of the rotor blades can withstand the temperature in the core of the engine and the outer portions can withstand the temperature of the freestream air. In general, the blades may be made of any suitable material capable of withstanding a temperature of up to 1300K, or up to 1620K, or a higher value that may correspond to a stagnation temperature at a higher flight Mach number. For example, the blades may be made of a ceramic matrix composite capable of withstanding a temperature of 1620K, meaning the blade is suitable for flight speeds of up to Mach 6. In general, the

[0057] The separation of the core flow and bypass flow so that only the core air flow is cooled by the heat exchanger 220 is also facilitated by the positioning of the bypass fan blades (the outer radial portion 270a of the rotor 270) at the rear of the engine, i.e. connected to the core turbine blades rather than the core compressor blades. This is particularly enabled by the arrangement of the fan blades being arranged behind the core inlet in which the pre-cooler is situated, so that no bypass air received by the fan blades is cooled by the heat exchanger beforehand.

[0058] Turning now to Figure 3, an alternative engine according to the present disclosure is illustrated. Similar to engine 200 depicted in Figure 2a or 2b, engine 300 of Figure 3 is also able to advantageously separate the core flow (first flow path 306) from the bypass flow (second flow path 308) so that only the core flow is cooled, meaning that the cooling capacity of the first heat exchanger 320 is not wasted on cooling the bypass flow, which would otherwise produce insignificant thrust benefits. The engine 300 depicted in Figure 3 is similar to engine 200 depicted in Figure 2a, and has corresponding components (e.g. combustion chamber 250 of engine 200 is similar, to identical to combustion chamber 350 of engine 300). It would be appreciated that whilst a core and bypass inlet arrangement according to that of Figure 2a is shown in Figure 3, some embodiments of an engine according to Figure 3 may instead include an arrangement of core and bypass inlets according to that shown in Figure 2b. Much like engine 200, engine 300 also comprises one or more rotors 370 (which correspond to one or more rotors 270) centred on the longitudinal axis. In the illustrated example, the engine comprises one rotor 370 that comprises an outer radial portion 370a and an inner radial portion 370b. Whilst the exemplary engines depicted in Figure 2a or 2b comprise the one or more rotors 270 arranged downstream of the turbomachinery components (including the combustion chamber 250 and turbine 260) such that the inner radial portion 270b operates as a turbine that is driven by the exhaust gases, the engine depicted in Figure 3 contrastingly has the one or more rotors 270 arranged upstream of the turbomachinery components (combustion chamber 350 and turbine(s) 360). In other words, the engine depicted in Figure 2a or 2b may be referred to as an "aft-fan" architecture, since the fan (i.e. the outer radial portions of the rotors 270) is positioned (along the longitudinal axis) behind or after the turbomachinery components of the engines core (i.e. the fan is positioned towards the back of the engine). The engine depicted in Figure 3 may be referred to as a "fore-fan" architecture, since the fan (i.e. the outer radial portions of the rotors 370) is positioned in front of the turbomachinery components and therefore towards the front of the engine.

[0059] As shown in Figure 3, and similar to rotors 270 shown in figure 2a or 2b, the outer radial portion 370a is arranged radially between the inner and outer casings, such that the outer radial portion 370a corresponds to the bypass flow received in the bypass inlet 312. The inner radial portion 370b is arranged radially inside of the inner casing 304, such that the inner radial portion 370b corresponds to the core flow receiving in the core inlet 310. The core flow in engine 300 is depicted by first flow path 306 that extends from the core inlet 310 to one or more exhaust nozzles via the heat exchanger heat exchanger 320, via the inner radial portion (370b) of the rotor 370, and the turbomachinery components in the core of the engine. Along the first flow path 306, the inner radial portion 370b is arranged so as to receive air downstream of the heat exchanger that has optionally been cooled by the heat exchanger (depending on the operating parameters of the engine). The cooled air received by the inner radial portion of the rotors 370 is then compressed by the rotors 370 and optionally by one or more additional compressor stages (not shown in the figure). For example, the one or more rotors 370 may function as low-pressure compressors that compress the cooled core flow by a first amount, and the engine may further comprise additional high-pressure compressors downstream of the rotors 370 to compress the already-compressed core air flow by a second amount. The turbomachinery components of the engine 300 comprise one of more turbines 360 each configured to drive rotation of a corresponding rotor 370. For example, engine 300 may comprise two turbine stages 360 and two rotors 370, wherein each turbine stage is connected to one of the rotors via a shaft (not shown) along the longitudinal axis of the engine. The engine may comprise additional turbine stages connected via respective shafts to corresponding high pressure compressor stages. During operation, the turbine stages are caused to rotate as a result of exhaust gas flowing flow the turbines, which in turn drives rotation of the rotors 370 via corresponding shafts. Thus, in this embodiment, the inner radial portion of the rotors function as compressor stages of the engine upstream of the combustion chamber, whereas the embodiment illustrated in Figure 2a or 2b has rotors with inner radial portions that function as turbine stages downstream of the combustion chamber.

[0060] As with the embodiment of Figure 2a or 2b, engine 300 also has a second flow path that extends from the bypass inlet 312 to one or more exhaust nozzles via the outer radial portion 370a of the one or more rotors 370. The second flow path bypasses the heat exchanger 320 and the turbomachinery components in the core of the engine. Along the second flow path 208, the outer radial portion 270a is arranged so as to receive the freestream bypass air flow that is received in the bypass inlet 212. As a result of the inner radial portion 270b of the rotor being driven by turbines 360, the outer radial portion 270a is driven to rotate since the outer radial portion and the inner radial portion of each rotor blade is connected (and are optionally integral, meaning they formed of the same component). Thus, in operation of the engine 300 at low flight Mach numbers, rotation of the outer radial portion 370a (and more generally, of the rotor 370 as a whole) causes the freestream bypass air flowing into the outer radial portion 370a to be compressed, thereby producing thrust in the manner that a conventional turbofan produces thrust. Thus, the outer radial portion of the rotor functions as a fan that compresses the bypass flow, whilst the inner radial portion functions as a compressor that compresses the core flow.

[0061] As mentioned above, rotor 370 may have a similar design to that depicted in Figure 4, wherein the rotor comprises a plurality of blades angularly spaced apart from one another and each arranged to extend from the centre of the rotor (at the longitudinal axis) outwards in a generally radial direction perpendicular to the longitudinal axis. The blades of the rotor 370, and in particular the inner radial portion of each blade may be shaped differently compared to the blades of rotor 270 for engine 200. This is because the inner radial portion of the rotor blades for engine 200 are designed to function as a turbine, since the rotor 270 is arranged downstream of the combustion chamber. In contrast, the inner radial portion of the rotor blades for engine 300 are designed to function as a compressor arranged upstream of the combustion chamber. Thus, it would be appreciated by the skilled person that the inner radial portions of the blades of rotors 270 and 370 may be shaped differently since they have different functions. With reference to Figures 2 and 3, it is noted that each of engines 200 and 300 may be comprise a single respective rotor 270, 370, or may comprise a plurality of rotors 270, 370. For example, as described in more detail below, the engines may comprise two adjacent rotors that are contrarotating. In other embodiments, either engine may comprise co-rotating rotors and intermediate stator vanes.

[0062] The architectures of each of engines 200 and 300 advantageously avoid cooling of the bypass flow. This is because for each engine, the heat exchanger 220 / 320 is arranged downstream of the core inlet 210 / 310. In other words, the heat exchangers are arranged inside the core inlet, whilst the bypass inlet is concentrically outside of the core inlet. The core inlet is defined by the inner casing 204 / 304, which physically separates the core and bypass flow. Because the heat exchanger is downstream of the core inlet and positioned inside the inner casing, it is physically separated from the bypass flow path. Put differently, the axial extent of the inner casing in the forward direction (i.e. the direction away from the turbomachinery components, or otherwise defined as the direction opposite to the flow of air when the engine is in operation) is greater than the axial extent of the first heat exchanger.

[0063] The operation of an engine according to Figure 2a, 2b or 3 in various flight speed regimes will now be described. Whilst reference is made to the components of Figure 2a in the following description, it will be appreciated that the same applies to the operation of the engine according to Figure 3. Reference to any components of engine 200 may also apply to the corresponding components of engine 300.

[0064] At low flight Mach numbers (for example from static to approximately Mach 2), the engine according to Figure 2a, 2b or 3 operates as a turbofan, with thrust produced by both the core flow exhaust gases being ejected from the engine and by the bypass flow being compressed by the outer radial portion 270a (i.e. fan portion) of the one or more rotors 270. As shown in Figure 2a, 2b or 3, the engine 200 further comprises a reheat system 290 configured to produce additional thrust by mixing fuel with the pressurised bypass flow and / or exhaust gases and burning the fuel / oxygen mixture in the exhaust section of the engine. The reheat system may be used to provide additional thrust at take-off, for example to provide the necessary acceleration required for lift-off. At low flight speeds, the benefits of pre-cooling the core airflow using the heat exchanger 220 may be negligible, and so in some examples the core air flow may bypass the heat exchanger and flow straight to the compressor 240. Examples of controlling the flow of air through a heat exchanger are described in GB2601367. This is because, at such low flight speeds, the stagnation temperature of the core air flow is not high enough to impact thrust performance of the engine core, or reach the temperature limits of the turbomachinery components within the core. As such the precooler may be selectively used depending on the flight speed and other parameters associated with the core turbomachinery components.

[0065] At higher flight Mach numbers (for example from Mach 2 to Mach 3), the pre-cooling of the core air flow becomes more effective and beneficial due to the increased temperature of the freestream air entering the core. As is known by the skilled person, pre-cooling of the core air flow enables the core to operate at higher-than-usual flight speeds. Thus, the engine of the present disclosure is capable of reaching higher flight speeds under turbofan / reheated turbofan operation than conventional, unprecooled engines, due to the advantages achieved by precooling freestream air before it enters the engine core. Thus, at such flight speeds, the engine is still capable of operating under turbofan operation in which thrust is produced by the core and bypass flows.

[0066] Beyond a flight speed of Mach 3 and approaching Mach 4 and higher, the freestream pressure seen in the bypass flow of the engine becomes high enough to achieve efficient ramjet operation (in which air received in the bypass inlet is naturally pressurised due in bypass duct due to the speed at which the air flows into the bypass). During operation of the engine 200 at flight speeds beyond Mach 3, the amount of fuel fed into the combustion chamber 250 in the core of the engine may be gradually reduced as the flight speed increases. This causes the thrust output of the core to be reduced, which in turn results in the thrust output provided by the outer radial portion of the rotor(s) 270 that pressurises the bypass flow to also reduce (since the outer radial portion of the rotors is driven directly by core exhaust gases flowing through the inner radial portion of the rotors). At the same time, whilst the fuel flow to the core and thus thrust output under turbofan operation is reduced, the fuel flow to the reheat system 290 is increased as the flight speed increases. This means that more thrust is produced by the engine under ramjet operation, in which the bypass air flow, that is pressurised both by natural ram pressure (due to very high flight speed) combined with pressurisation by the outer radial portion 270a of the rotor, is mixed with fuel and burnt by the reheat system 290 to produce thrust.

[0067] The engine's core can continue to be "throttled back" (i.e. the thrust produced from the core can continue to be reduced by reducing the fuel flow to the core) to the point where the core is shut down and produces no thrust, since at this point (e.g. Mach 4 and above the ramjet alone is producing sufficient thrust to maintain or accelerate the flight speed. When the core is shut down, it may be allowed to windmill as air flows through the core. This means that, although the core is producing no thrust, the turbomachinery components of the core may be allowed to rotate as they are driven by the high-speed airflow through the core. When the core is in windmill mode, the flow of air through the core is considerably lower than when the core is producing thrust. This is particularly advantageous since the reduced flow through the core means that less cooling budget is used in the heat exchanger, meaning the cooling budget (i.e. the heat capacity of the coolant, which may be the engine's fuel) can be used to cool other onboard systems. A further advantage of allowing the core to windmill is that power can be extracted from the rotating turbomachinery components in the core. In other words, the high-speed air flow through the core drives rotation of the turbomachinery components in the core which in turn can be used to generate electrical power using conventional electrical generation and storage systems.

[0068] In other examples, the core of the engine may be locked (preventing windmilling) when the engine is operating solely as a ramjet (no thrust being produced in the core). This may be achieved by applying a large torque to the turbomachinery components in the core or by otherwise applying braking to the core. An advantage of locking the core is that airflow through the core is greatly reduced compared to allowing the core to windmill, which further reduces the cooling budget used in the core. Locking the core may result in 70% less airflow through the core compared to windmilling, meaning that 70% less cooling budget is used if the core is locked.

[0069] In yet other examples, air flow through the core may be shut off entirely (i.e. the air flow into the engine is prevented from flowing through the core, and may otherwise be diverted to the bypass). Shutting off flow from the core frees up cooling budget even further, which can then be used elsewhere in the engine or aircraft more generally.

[0070] Generally speaking, the engine according to Figure 2a or 2b therefore provides a solution for transitioning between a turbofan regime and a ramjet regime, in which there is a smooth transition that does not require a discrete, step-like change between an engine operating as a turbofan and a ramjet, nor does the engine require separate intakes and / or exhausts for each of the turbofan and ramjet. Instead, the turbofan and ramjet components are combined into a single engine architecture in which a gradual and smooth transition between turbofan and ramjet operation can be achieved. The speeds at which this transition occurs is low enough that sufficient thrust can be produced by the core flow and pressurised bypass flow, but also high enough that sufficient pressure is generated in the bypass flow to produce sufficient thrust under ramjet operation. This is specifically achieved as a result of at least: 1) the heat exchanger arranged to cool the air flowing into the engine core, which allows the engine to operate at higher-than usual speeds as a turbofan (thrust produced by the core and the pressurised bypass flow); and 2) the outer radial portion 270a of the rotor, which pressurises the bypass flow and in turn increases the air pressure in the exhaust (where the reheat system is situated), effectively allowing the engine to efficiently operate at lower-than-usual speeds as a ramjet (in other words, the pressure of the bypass air in the exhaust is increased by the additional pressurisation provided by the fan). By increasing the speed at which the turbofan can operate efficiently, and reducing the speed at which the ramjet can operate efficiently, an overlap in flight speeds between the two engine regimes can be achieved, so that a gradual transition from one regime to the other can be carried out without a significant loss of thrust during the transition.

[0071] A further advantage achieved by an engine according to Figure 2a or 2b is that at high flight Mach numbers, the core of the engine has a very low-pressure ratio across it. This is because, at high flight Mach numbers, the vast majority of air flowing into the engine flows through the bypass to the exhaust (since the engine is operating as a ramjet, and thrust produced by the core is reduced). In comparison, only a small portion of the total air entering the engine flows into the core of the engine, since at such high flight Mach numbers (i.e. beyond Mach 3) the fuel supplied to the core and hence thrust produced by the core is reduced. This effectively "back-pressures" the core of the engine, which further reduces the amount of air that flows through the core. Less air flow into the core reduces the core temperature, meaning that less cooling is required by the heat exchanger 220. In other words, the heat load into the working fluid of the heat exchanger (e.g. the engine's fuel), is reduced, effectively freeing up the heat capacity of the working fluid to be used elsewhere in the engine or aircraft.

[0072] Beyond the flight Mach numbers at which the engine transitions between operating as a turbofan and operating as a ramjet (e.g. Mach 3 to Mach 4), the engine may operate solely as a ramjet, with all of the thrust being produced by the bypass flow mixing with fuel and being combusted in the exhaust by the reheat system 290. Under such operation, the rotor is not required to pressurise the bypass flow for ramjet operation since the Mach numbers are already high enough to produce sufficient pressure in the exhaust. This means that the outer radial portion 270a of the rotor 270 "windmills" in the bypass flow. In other words, the flow of the air through the bypass duct drives rotation of the outer-radial portion of the rotor, in turn driving rotation of the rotor as a whole (this is in contrast to turbofan operation, where rotation of the rotor pressurises the bypass flow). At such flight speeds, it is desirable for the outer-radial portions of the rotor, which are positioned in the bypass flow, to provide zero or minimal blockage to the bypass flow. Otherwise, any blockage caused by the outer radial portions of the rotor will result in a pressure loss, leading to a loss of thrust produced by the ramjet or less efficient ramjet operation. To minimise blockage caused by the rotor in the bypass flow at such high flight speeds (i.e. Mach 4 and above) some embodiments of the present disclosure comprise two contra-rotating rotors, meaning that whilst one rotor is arranged to rotate in a first direction (e.g. clockwise) about the longitudinal axis, a second rotor is arranged to rotate in the opposite direction (e.g. counter-clockwise) about the longitudinal axis. In some embodiments, the two rotors are adjacent to one another along the longitudinal axis, meaning there are no stator vanes (e.g. static, non-rotating blades angularly spaced apart from one another and also arranged to extend from the longitudinal axis outwards in a radial direction perpendicular to the longitudinal axis) positioned between the rotors along the longitudinal direction. Such an arrangement of adjacent contra-rotating rotors without intermediate stators is referred to herein as a stator-less contra-rotating rotor arrangement, and is particularly advantageous at preventing compressor choke and minimising pressure loss due to blockage of the bypass flow. In other words, the stator-less contra-rotating rotor arrangement provides minimal blockage to the bypass flow compared to conventional compressor fan arrangements that include multiple rotor and stator stages.

[0073] The use of a stator-less contra-rotating rotor arrangement allows the rotational speed of the rotors to be selected such that all the diffusion that usually requires two co-rotating rotors (rotating in the same direction) and two stator rows in a conventional fan / compressor can be achieved within the same aerodynamic limitations of diffusion in 2 blade rows. In other words, the necessary deceleration of the bypass flow that is required for pressurisation of the bypass flow that is achieved using a conventional arrangement (two co-rotating rotors and two stators) can be achieved using two contra-rotating rotors. This in effect allows the fan to be designed with a low exit flow Mach numbers such that when the system goes into high Mach ramjet operation the Mach numbers are kept well below choke in the blade rows and the bypass, mixer and ramjet duct. In other words, even when the system is operating in the high Mach Ramjet regime, with correspondingly high mass flow rates through the bypass, the contra-rotating stator-less fan blades enable the velocity of the flow through the blades to remain below Mach 1, thereby preventing compressor choke which would otherwise limit the mass flow rate through the blades. This is compounded with the advantage of minimal-blockage due to absence of stator vanes, meaning that overall, outer radial portion 270a of the rotor 270 being arranged within the bypass flow has little impedance on the operation of the ramjet at high flight Mach numbers, due to minimal pressure loss in the bypass.

[0074] A further advantage of using contra-rotating stator-less fan blades in the disclosed engine arrangements is that it is possible to achieve good levels of power extraction from the fan blades when they are "windmilling" at high flight Mach numbers (e.g. Mach 4 and above). Power extraction from the windmilling fan can be carried out using conventional electrical generation and storage means by generating electrical power from the rotating fan (rotors 270). Since the stator-less contrarotating fan blade arrangement can prevent choke even at high flight speeds, it is possible to generate electrical power from the rotation of these fans without risking engine choke. This is in comparison to co-rotating fan blade arrangements with intermittent stators, which are more likely to choke at lower speeds and are therefore less effective at generating electrical power. Therefore, in some embodiments, it is possible to generate electrical power from the rotating (windmilling) rotor blades when the engine is producing thrust solely under ramjet operation.

[0075] It should be understood that one or more of the aforementioned embodiments may be combined with one another. It should also be understood that the disclosure includes an aircraft, flying machine, or aerospace vehicle comprising any one or more of the exemplary engines as described above.

[0076] Various modifications may be made to the described embodiment(s) without departing from the scope of the invention as defined by the accompanying claims.

Claims

CLAIMS1. An engine, comprising: a core air inlet configured to receive air and generally centred on a longitudinal axis of the engine; a bypass air inlet configured to receive air; a first heat exchanger arrangement arranged downstream of the core inlet and configured to cool air received in the core inlet; one or more turbomachinery components configured to receive air and arranged downstream of the first heat exchanger arrangement, and centred on the longitudinal axis; at least one rotor centred on the longitudinal axis and having a radial direction perpendicular to the longitudinal axis; a first flow path arranged to extend through the core inlet, the first heat exchanger arrangement, the one or more turbomachinery components and an inner radial portion of the at least one rotor; and a second flow path arranged to extend through the bypass inlet and an outer radial portion of the at least one rotor, the second flow path arranged to bypass the first heat exchanger arrangement, and the one or more turbomachinery components.

2. An engine as claimed in claim 1, wherein the at least one rotor is arranged downstream of the one or more turbomachinery components.

3. An engine as claimed in claim 1, wherein the at least one rotor is arranged upstream of the one or more turbomachinery components and downstream of the first heat exchanger.

4. An engine as claimed in any preceding claim, wherein the core inlet is defined by a generally cylindrical inner casing within which the first heat exchanger arrangement is arranged, and the bypass inlet defined by a generally cylindrical outer casing arranged concentric with and radially outside the inner casing, and wherein the first flow path extends through the inner casing and the second flow path extends between the inner and outer casings.

5. An engine as claimed in claim 4, wherein the axial extent of the inner casing in a direction away from the turbomachinery components is greater than an axial extent of the first heat exchanger arrangement in the direction away from the turbomachinery components.

6. An engine as claimed in any preceding claim, further comprising a main inlet upstream of the core inlet and bypass inlet, wherein the first and second flow paths extend from a common flow path that extends through the main inlet.

7. An engine as claimed in any preceding claim, wherein the at least one rotor comprises a plurality of blades angularly spaced apart from each other and each arranged to extend generally along the radial direction of the rotor, wherein each of the blades has a root corresponding to the inner radial portion of the rotor and a tip corresponding to the outer radial portion of the rotor, the root being arranged closer to the longitudinal axis than the tip such that the first flow path is arranged to extend adjacent the roots of the blades and the second flow path is arranged to extend adjacent the tips of the blades.

8. An engine as claimed in claim 2 or any of claims 4 to 7 when dependent on claim 2, wherein the one or more turbomachinery components comprise at least one compressor downstream of the first heat exchanger arrangement and at least one combustion chamber.

9. An engine as claimed in claim 3 or any of claims 4 to 7 when dependent on claim 3, wherein the one or more turbomachinery components comprise at least one combustion chamber and at least one turbine downstream of the at least one combustion chamber, wherein the at least one turbine is coupled to a respective one of the at least one rotors via a shaft.

10. An engine as claimed in claim 8 or 9, wherein the at least one combustion chamber comprises an air-breathing combustion chamber configured to receive compressed air for combustion, wherein the combustion chamber is configured to mix the received compressed air with fuel, and combust the compressed air and fuel mixture to produce exhaust gas.

11. An engine as claimed in claim 10 when dependent on claim 8, wherein the at least one rotor is arranged downstream of the turbine such that: the inner radial portion of the at least one rotor is arranged to receive the exhaust gas to thereby cause rotation of the rotor about the longitudinal axis as a result of the exhaust gas flowing through the inner radial portion of the rotor; androtation of the rotor caused by the exhaust gas flowing through the inner radial portion of the rotor causes the outer portion of rotor to compress air received by the bypass inlet flowing along the second flow path.

12. An engine as claimed in claim 7 or any of claims 8 to 11 when dependent on claim 7 , wherein the root and tip of each blade are made of the same material.

13. An engine as claimed in claim 10 when dependent on claim 9, wherein the at least one turbine is arranged to rotate as a result of the exhaust gas flowing through the at least one turbine, and wherein rotation of the at least one turbine in turn causes rotation of the respective at least one rotor via the respective shaft; and rotation of the rotor caused by rotation of the turbine causes the inner radial portion of the rotor to compress air receiving in the core inlet along the first flow path, and causes the outer portion of rotor to compress air received by the bypass inlet flowing along the second flow path.

14. An engine as claimed in any preceding claim, wherein the at least one rotor comprises two rotors each arranged along and centred on the longitudinal axis of the engine and arranged to rotate independently about the longitudinal axis.

15. An engine as claimed in claim 14, wherein the two rotors are arranged to rotate in opposite directions.

16. An engine as claimed in claim 14 or 15, wherein the two rotors are adjacent along the longitudinal axis and wherein the engine does not comprise stator vanes positioned in between the two rotors along the longitudinal direction.

17. An engine as claimed in any preceding claim, further comprising an exhaust nozzle, wherein the first and second flow paths are arranged to flow from their respective air inlets to the exhaust nozzle.

18. An engine as claimed in claim 17, further comprising a reheat system arranged downstream of the at least one rotor, wherein the first and second flow paths both extend through the reheat system.

19. A method of operating an engine as claimed in claim 1, the method comprising: causing a first portion of air to flow along the first flow path such that at least a portion of the air received by the core inlet is cooled in the first heat exchanger arrangement; causing, using the one or more turbomachinery components, combustion of the first portion of air to produce a first exhaust gas; causing at least a portion of the first exhaust gas to flow through an inner radial portion of the at least one rotor, thereby driving rotation of the rotor; causing the first exhaust gas to exit the engine through an exhaust nozzle to produce a first amount of thrust; causing a second portion of air, different from the first portion of air, to flow along the second flow path such that at least a portion of the air received by the bypass inlet is configured to extend through the outer portion of the one or more rotors; compressing, using the outer portion of the of the one or more rotors, at least a portion of the air received by the bypass inlet; combusting, using a reheat system, the compressed air to produce a second exhaust gas; and causing the second exhaust gas to exit the engine through the exhaust nozzle to produce a second amount of thrust.

20. The method of claim 19, wherein, when a flight speed of the engine reaches a first threshold speed, the method comprises reducing an amount of fuel supplied to the one or more turbomachinery components to reduce the first amount of thrust produced by the engine.

21. The method of claim 20, wherein when a flight speed of the engine reaches a second threshold speed that is greater than the first threshold speed, the method comprises stopping fuel flow to the one or more turbomachinery components such that the first amount of thrust becomes zero.

22. The method of claim 21, wherein when the flight speed of the engine is greater than or equal to the second threshold speed, the first portion of air flowing along the first flow path drives rotation of the one or more turbomachinery components, and wherein the method further comprises generating electrical power using the rotation of the one or more turbomachinery components.

23. The method of claim 21, wherein when the flight speed of the engine is greater than or equal to the second threshold speed, the method comprises preventing rotation of the one or moreturbomachinery components to reduce the amount of the first portion of air flowing along the first flow path.

24. The method of any of claims 21 to 23, wherein when the flight speed of the engine is greater than or equal to the second threshold speed, the second portion of air flowing along the second flow path drives rotation of the one or more rotors, and wherein the method further comprises generating electrical power using the rotation of the one or more rotors.

25. An aircraft, flying machine, or aerospace vehicle comprising an engine as claimed in any of claims 1 to 18.

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