Aircraft with a single fluid inlet aperture

Abstract

An aircraft comprises a machine body. The machine body encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a combustor module, a turbine module, and an exhaust module. The machine body comprises a single fluid inlet aperture, with the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the heat exchanger module. The heat exchanger module is configured to transfer a waste heat load from the gas turbine engine and the ancillary systems to the fluid cooling flow prior to an entry of the entire fluid cooling flow into the fan module.

Claims

1. An aircraft comprising a machine body, the machine body enclosing a turbofan gas turbine engine and a plurality of ancillary systems, the turbofan gas turbine engine comprising, in axial flow sequence, a heat exchanger module, a fan module, a compressor module, a combustor module, a turbine module, and an exhaust module; and wherein the machine body comprises a single fluid inlet aperture, the fluid inlet aperture being configured to allow a fluid cooling flow to enter the machine body and to pass through the heat exchanger module, the heat exchanger module being configured to transfer a waste heat load from the gas turbine engine and the ancillary systems to the fluid cooling flow prior to an entry of the entire fluid cooling flow into the fan module, the heat exchanger module including a plurality of heat exchanger elements that each extends from a portion along a center axis of the heat exchanger module to an outer circumferential portion of the heat exchanger module, so that the plurality of heat exchanger elements are arranged as a circumferential array of radially extending vanes that extend across an entirety of an inlet opening of the heat exchanger module; the fan module includes a plurality of fan blades mounted on a fan hub; and the heat exchanger module further includes a heat exchanger hub on which the plurality of radially extending vanes are mounted, wherein the heat exchanger hub is separate and spaced apart from the fan hub along the center axis.

2. The aircraft as claimed in claim 1, wherein a proportion B.sub.COMB of the fluid cooling flow passes sequentially through the compressor, combustor, and turbine modules, wherein the B.sub.COMB parameter is defined in the range of 0.20 to 0.71.

3. The aircraft as claimed in claim 2, wherein the B.sub.COMB parameter is defined in the range of 0.29 to 0.71.

4. The aircraft as claimed in claim 1, the machine body further comprising a single fluid exhaust aperture, and wherein the single fluid exhaust aperture is configured to channel the fluid flow from the exhaust module out of the machine body.

5. The aircraft as claimed in claim 1, wherein the fan module comprises a plurality of fan blades defining a fan diameter (D), and the fan diameter D is within the range of 0.3 m to 2.0 m.

6. A method of operating an aircraft, the aircraft comprising a machine body, the machine body enclosing a turbofan gas turbine engine and a plurality of ancillary systems, the turbofan gas turbine engine comprising, in axial flow sequence, a heat exchanger module, a fan module, a compressor module, a combustor module, a turbine module, and an exhaust module; and wherein the method comprises the steps of: (i) providing the machine body; (ii) arranging the fan module, the compressor module, the combustor module, the turbine module, and the exhaust module within the machine body; (iii) providing the machine body with a single fluid inlet aperture, the fluid inlet aperture being configured to allow a fluid flow to enter the machine body and to pass through the heat exchanger module; (iv) configuring the heat exchanger module to transfer a waste heat load from the gas turbine engine and the ancillary systems to the fluid flow prior to an entry of the fluid flow into the fan module, by providing the heat exchanger module to include a plurality of heat exchanger elements that each extends from a portion along a center axis of the heat exchanger module to an outer circumferential portion of the heat exchanger module so that the plurality of heat exchanger elements are arranged as a circumferential array of radially extending vanes that extend across an entirety of an inlet opening of the heat exchanger module, the heat exchanger module further including a heat exchanger hub on which the plurality of radially extending vanes are mounted, wherein the fan module includes a plurality of fan blades mounted on a fan hub and wherein the heat exchanger hub is separate and spaced apart from the fan hub along the center axis; and (v) operating the engine such that the entire fluid flow enters the fan module.

7. The method as claimed in claim 6, wherein step (v) comprises the step of: (v)' operating the engine such that the entire fluid flow enters the fan module, and a proportion B.sub.COMB of the fluid flow passes sequentially through the compressor, combustor, and turbine modules, and the B.sub.COMB parameter is defined in the range of 0.20 to 0.71.

8. The method as claimed in claim 6, wherein step (iii) comprises the additional following step of: (iii)' providing the machine body with a single fluid exhaust aperture, the single fluid exhaust aperture being configured to channel the fluid flow from the exhaust module out of the machine body.

9. The aircraft as claimed in claim 5, wherein the fan diameter D is within the range of 0.4 m to 1.5 m.

10. The aircraft as claimed in claim 5, wherein the fan diameter D is within the range of 0.7 m to 1.0 m.

11. The aircraft as claimed in claim 1, wherein the heat exchanger module has a flow diameter that is greater than a fan diameter of fan blades of the fan module.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) There now follows a description of an embodiment of the disclosure, by way of non-limiting example, with reference being made to the accompanying drawings in which:

(2) FIG. 1 shows a schematic perspective view of an aircraft according to a first embodiment of the disclosure;

(3) FIG. 2 shows a schematic sectional view of the aircraft of FIG. 1;

(4) FIG. 3 shows a schematic view of the heat exchanger module and fan module of the turbofan gas turbine engine of the aircraft of FIG. 1;

(5) FIG. 4 shows a schematic sectional view of the heat exchanger module and fan module of the turbofan gas turbine engine of the aircraft of FIG. 1; and

(6) FIG. 5 shows an alternative arrangement of the heat exchanger module and fan module shown in FIG. 4.

(7) It is noted that the drawings may not be to scale. The drawings are intended to depict only typical aspects of the disclosure, and therefore should not be considered as limiting the scope of the disclosure. In the drawings, like numbering represents like elements between the drawings.

DETAILED DESCRIPTION

(8) Referring to FIGS. 1 and 2, an aircraft according to a first embodiment of the disclosure is designated generally by the reference numeral 100. The aircraft 100 comprises a machine body 102 in the form of a fuselage with wings and a tail plane. The machine body 102 encloses a turbofan gas turbine engine 110, together with a plurality of ancillary systems 104.

(9) Within the machine body 102 there is a cockpit volume 106, a payload volume 108, and a plurality of ancillary systems 104.

(10) The turbofan gas turbine engine 110 comprises, in axial flow sequence, a heat exchanger module 120, a fan module 130, a compressor module 140, a combustor module 150, a turbine module 160, and an exhaust module 170. The turbofan gas turbine engine 110 further comprises a second heat transfer element 124. The second heat transfer element 124 takes the form of a heat exchanger that uses the engine's fuel as a cooling medium.

(11) The fan module 130, compressor module 140, combustor module 150, turbine module 160, and exhaust module 170, together forming the core engine, are enclosed within an outer casing 180. An annular bypass duct 182 is defined between the core engine and the outer casing 180.

(12) The heat exchanger module 120 comprises a plurality of first heat exchanger elements 122. In the present arrangement, illustrated in FIG. 3, the first heat exchanger elements 122 are arranged as a circumferential array of radially extending vanes 122. The inlet air flow 101 passes over the surface of the first heat exchanger elements 122 as the air flow passes through the heat exchanger module 120.

(13) The heat exchanger module 120 has a total heat rejection capacity. The total heat rejection capacity is the amount of waste heat energy that can be dissipated into an air flow passing through the heat exchanger module 120.

(14) The machine body 102 comprises only one fluid inlet aperture 112. The fluid inlet aperture 112 is configured to allow an intake air flow 101 to enter the machine body 102. In other words, there is only one inlet aperture 112 on the machine body 102 through which an air flow 101 can enter the machine body 102.

(15) The intake air flow 101 passes through the heat exchanger module 120 and subsequently passes through the fan module 130. Once through the fan module, the air flow divides into a first flow (not shown) and a second flow (not shown). The first flow (the ‘core’ flow) passes sequentially through the core engine, i.e. sequentially through compressor module 140, the combustor module 150, the turbine module 160, and the exhaust module 170. The second flow (the ‘bypass’ flow) exits the fan module 130 and passes through the annular bypass duct 182 to the exhaust module 170.

(16) The machine body 102 further comprises only one fluid exhaust aperture 104. The air flow from the exhaust module 170 exits the machine body 102 through the single fluid exhaust aperture 104. In other words, there is only one exhaust aperture 104 in the machine body 102 through which an air flow 101 can exit the machine body 102.

(17) As outlined above, the machine body 102 of present disclosure includes only two apertures 112,114 in its outer surface; an inlet aperture 112 allowing an air flow into the machine body and an exhaust aperture 114 allowing the air flow to exhaust from the machine body. The presence of apertures in the machine body 102 causes parasitic aerodynamic drag on the machine body 102. As in the present arrangement, the use of only two apertures 112,114 in the machine body 102 minimises this parasitic aerodynamic drag.

(18) In the present arrangement, the fan assembly 130 comprises two fan stages (not shown), with each fan stage comprising a plurality of fan blades (not shown). In the present arrangement each fan stage has the same fan diameter 132, with the respective plurality of fan blades defining a fan diameter of 0.9 m. In an alternative arrangement, the two fan stages may have different fan diameters 132 each defined by the corresponding plurality of fan blades. As previously mentioned, the fan diameter (D) 132 is defined by a circle circumscribed by the leading edges of the respective plurality of fan blades.

(19) In use, both the turbofan gas turbine engine 110 and the ancillary systems 104 generate waste heat energy that is required to be dissipated to ensure the safe operation of the turbofan engine 110 and the ancillary systems 104.

(20) As shown in FIG. 3, the heat exchanger module 120 has a flow area (A.sub.HEX) 126. The heat exchanger module flow area 126 is the cross-sectional area of the heat exchanger module 120 through which an inlet air flow 101 passes before being ingested by the fan module 130. In the present arrangement, the heat exchanger module flow area 126 has an annular cross-section and corresponds directly to the shape of the air flow passing through the heat exchanger module 120.

(21) The fan module 130 has a corresponding flow area (A.sub.FAN) 134. The fan module flow area 134 is the cross-sectional area of the fan module 130 through which an inlet air flow 101 passes before separating into the core engine flow and the bypass flow. The fan assembly flow area 134 has an annular shape since it corresponds to the annular area swept by the fan blades 131.

(22) In the present arrangement (illustrated in FIG. 4) the heat exchanger module flow area 126 is equal to the fan module flow area 134, and the corresponding ratio of A.sub.HEX/A.sub.FAN is equal to 1.0.

(23) The heat exchanger module 120 has a flow diameter (E) 128, which is the diameter of the air flow passing through the heat exchanger module 120. In the present arrangement, shown in FIG. 4, the heat exchanger module flow diameter 128 is equal to the fan diameter 132. In an alternative arrangement (see FIG. 5) the heat exchanger module flow diameter 228 is greater than the fan diameter 132.

(24) The heat exchanger module 120 is configured to transfer a waste heat load from the gas turbine engine 110 and the ancillary systems 104 to the fluid flow 101 prior to the entry of the fluid flow 101 into the fan module 130. A first fluid 116, which in this embodiment is a synthetic oil, is circulated through hot portions of the turbofan engine 110 and the ancillary systems 104 to collect waste heat energy.

(25) As outlined above, once the air flow 101 has passed through the fan module 130, the air flow 101 divides into two flow portions, a first portion being the so-called ‘core flow’, and a second portion being the so-called ‘bypass flow’. The core flow enters the compressor module 140 and continues sequentially through the combustor module 150, turbine module 160, and exhaust module 170. The bypass flow passes through the annular bypass duct 182 into the exhaust module 170. The core flow and the bypass flow join in the exhaust module 170 and are exhausted from the machine body 102 through the exhaust aperture 114.

(26) The core flow can be characterised by the parameter B.sub.COMB which represents the proportion of the fluid flow 101 entering the machine body 102 that subsequently passes sequentially through the compressor, combustor, turbine and exhaust modules 140,150,160,170. In the present arrangement, the turbofan engine 110 has a bypass ratio of 2. In this arrangement, the turbofan engine 110 can be characterised by a B.sub.COMB parameter of 0.29.

(27) In use, at a cruise condition the aircraft 100 is capable of maintaining a sustained airspeed V (in metres per second, m/s). At this sustained airspeed V, the heat exchanger module 120 transfers a total waste heat energy load H (in Watts, W) to the fluid flow 101.

(28) In use, when the sustained airspeed V of the aircraft 100 is less than Mach 1.0 (i.e. the aircraft 100 is in subsonic flight) the first fluid 116 is circulated through the first heat exchanger elements 122 to dissipate the waste heat energy contained in the first fluid 116 to the inlet fluid flow 101.

(29) When the sustained airspeed V of the aircraft 100 exceeds, for example, Mach 1.0 (i.e. supersonic flight conditions) the temperature of the inlet fluid flow (T.sub.A) 101 increases. This temperature increase will significantly reduce the efficiency of the transfer of the waste heat energy from the first fluid 116 to the inlet fluid flow 101. Continued increase in the sustained airspeed V of the aircraft 100 will cause a continued rise in the temperature of the inlet fluid flow 101. Once this temperature T.sub.A reaches the temperature of the first fluid (T.sub.F) 116 it will not be possible to dissipate waste heat energy to the inlet fluid flow 101 via the first heat transfer elements 122.

(30) Consequently, in the arrangement of the present invention, when the airflow temperature T.sub.A is equal to or greater than the first fluid temperature T.sub.F, the flow of the first fluid 116 is routed through the second heat transfer element 124. The second heat transfer element 124 uses the fuel supply to the turbofan engine 110 as the cooling medium.

(31) While the temperature of the inlet fluid flow T.sub.A 101 will increase at and above a sustained airspeed of, for example, M1.0, the temperature of the engine's fuel will remain substantially constant. By routing the first fluid 116 through the second heat transfer element 124 it becomes possible to continue to dissipate the waste heat energy from the turbofan engine 110 and the ancillary systems 104 even when the temperature of the inlet fluid flow 101 is greater than the temperature of the first fluid 116.

(32) Note that the terms “low-pressure turbine” and “low-pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine. In some literature, the “low-pressure turbine” and “low-pressure compressor” referred to herein may alternatively be known as the “intermediate-pressure turbine” and “intermediate-pressure compressor”. Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

(33) Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example.

(34) The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

(35) It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

(36) The invention includes methods that may be performed using the subject devices. The methods may comprise the act of providing such a suitable device. Such provision may be performed by the end user. In other words, the “providing” act merely requires the end user obtain, access, approach, position, set-up, activate, power-up or otherwise act to provide the requisite device in the subject method. Methods recited herein may be carried out in any order of the recited events which is logically possible, as well as in the recited order of events.

(37) In addition, where a range of values is provided, it is understood that every intervening value, between the upper and lower limit of that range and any other stated or intervening value in that stated range, is encompassed within the invention.

(38) Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.