Turbofan engine with heat exchanger module having optimized fan to element area parameter

Abstract

A turbofan gas turbine engine includes, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, and a turbine module. The fan assembly includes fan blades defining a corresponding fan area (A.sub.FAN). The heat exchanger module is in fluid communication with the fan assembly by an inlet duct, and includes radially-extending vanes arranged in a circumferential array with at least one vane including a heat transfer element for heat transfer from a first fluid contained within each element to an airflow passing over a surface of each heat transfer element before entering the fan assembly inlet. Each heat transfer element extends axially along the corresponding vane, with a swept heat transfer element area (A.sub.HTE) being the wetted surface area of all heat transfer elements in contact with the airflow. A Fan to Element Area parameter F.sub.EA of A.sub.HTE/A.sub.FAN lies in the range of 47 to 132.

Claims

1. A turbofan gas turbine engine comprising, in axial flow sequence: a heat exchanger module comprising a plurality of radially-extending hollow vanes arranged in a circumferential array, at least one of the vanes comprising a heat transfer element positioned within the vane and that extends axially along the vane and is for transfer of heat from a first fluid contained within the heat transfer element to an airflow passing through a hollow portion of the vane and over a surface of the heat transfer element prior to entry of the airflow into an inlet to a fan assembly; the fan assembly, which (i) is in fluid communication with the heat exchanger module by way of an inlet duct and (ii) comprises a plurality of fan blades defining a corresponding fan area (A.sub.FAN); a compressor module; and a turbine module, wherein a swept heat transfer element area (A.sub.HTE) is a wetted surface area of all of the heat transfer elements in contact with the airflow, a Fan to Element Area F.sub.EA parameter is defined as: $F_{\mathrm{EA}}=\frac{A_{\mathrm{HTE}}}{A_{\mathrm{FAN}}}$ and the F.sub.EA parameter lies in a range of 47 to 132.

2. The turbofan gas turbine engine as claimed in claim 1, wherein the F.sub.EA parameter lies in a range of 57 to 97.

3. The turbofan gas turbine engine as claimed in claim 1, wherein the F.sub.EA parameter lies in a range of 76 to 123.

4. The turbofan gas turbine engine as claimed in claim 1, wherein the fan assembly comprises a plurality of fan blades defining a fan diameter D, and the fan diameter D is within a range of 0.3 m to 2.0 m.

5. The turbofan gas turbine engine as claimed in claim 1, wherein the heat exchanger module has a flow area A.sub.HEX, the fan assembly has a flow area A.sub.FAN, and a ratio of A.sub.FAN to A.sub.HEX is in a range of 0.3 to 0.8.

6. The turbofan gas turbine engine as claimed in claim 1, wherein the heat exchanger module has a fluid path diameter E, and the fluid path diameter E is greater than a fan diameter D.

7. The turbofan gas turbine engine as claimed in claim 1, further comprising an outer housing enclosing the heat exchanger module, the fan assembly, the compressor module, and the turbine module, wherein an annular bypass duct is defined between the outer housing and an engine core comprising the compressor and turbine modules, a bypass ratio is defined as a ratio of a mass air flow rate through the bypass duct to a mass air flow rate through the engine core at cruise conditions, and the bypass ratio is less than 2.0.

8. The turbofan gas turbine engine as claimed in claim 1, wherein the fan assembly has two or more fan stages, at least one of the fan stages comprising a plurality of fan blades defining a fan diameter D.

9. The turbofan gas turbine engine as claimed in claim 1, wherein the at least one vane comprises a plurality of the heat transfer elements.

10. A method of operating an aircraft comprising the turbofan gas turbine engine as claimed in claim 1, the method comprising taking off from a runway, wherein a maximum rotational speed of the turbine during take-off is in a range of from 8500 rpm to 12500 rpm.

11. A method of operating a turbofan gas turbine engine, the turbofan gas turbine engine comprising, in axial flow sequence, (i) a heat exchanger module, (ii) an inlet duct, (iii) a fan assembly comprising a plurality of fan blades defining a corresponding fan area (A.sub.FAN), (iv) a compressor module, and (v) a turbine module, and the method comprising the steps of: (i) providing the fan assembly, the compressor module, and the turbine module; (ii) positioning the heat exchanger module in fluid communication with the fan assembly by way of the inlet duct; (iii) providing the heat exchanger module with a plurality of radially-extending hollow vanes arranged in a circumferential array, at least one of the vanes comprising a heat transfer element positioned within the vane and extending axially along the vane, a swept heat transfer element area (A.sub.HTE) being a wetted surface area of all the heat transfer elements in contact with an airflow passing over a surface of each heat transfer element prior to entry of the airflow into an inlet to the fan assembly, a Fan to Element Area F.sub.EA parameter being defined as: $F_{\mathrm{EA}}=\frac{A_{\mathrm{HTE}}}{A_{\mathrm{FAN}}}$ and the F.sub.EA parameter lying in a range of 47 to 132; and (iv) operating the engine such that the airflow passes through a hollow portion of the vane and transfers heat energy from a first fluid contained within each heat transfer element to the airflow.

12. The method of claim 11, wherein the F.sub.EA parameter of step (iii) lies in a range of 57 to 97.

13. The method of claim 11, wherein the F.sub.EA parameter of step (iii) lies in a range of 76 to 123.

14. The method of claim 11, wherein the at least one vane comprises a plurality of the heat transfer elements.

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 part-sectional view of a turbofan gas turbine engine according to the prior art;

(3) FIG. 2 shows a schematic part-sectional view of a turbofan gas turbine engine according to a first embodiment of the disclosure;

(4) FIG. 3 shows a schematic part-sectional view of a turbofan gas turbine engine according to a second embodiment of the disclosure;

(5) FIG. 4 shows a perspective schematic view of the heat exchanger module of the turbofan engine of FIG. 2 showing the circumferential array of vanes forming the heat exchanger module; and

(6) FIG. 5 shows a perspective schematic part-sectional view of one of the vanes of the heat exchanger module of 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) FIG. 1 illustrates a conventional turbofan gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a two-stage propulsive fan 13 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, an intermediate-pressure turbine 18, a low-pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 13 is attached to and driven by the low-pressure turbine 19 via a shaft 26.

(9) In use, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed into the high-pressure compressor 15 where further compression takes place. The compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high-pressure, intermediate-pressure, and low-pressure turbines 17, 18, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 by a suitable interconnecting shaft 27. The low-pressure compressor 14 drives the intermediate-pressure turbine 18 via a shaft 28.

(10) 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 13) 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 13 may be referred to as a first, or lowest pressure, compression stage.

(11) Other turbofan 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 fans and/or compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 23 meaning that the flow through the bypass duct 22 has its own nozzle 23 that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core engine 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. 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.

(12) The geometry of the turbofan 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.

(13) Referring to FIG. 2, a turbofan gas turbine engine according to a first embodiment of the disclosure is designated generally by the reference numeral 100. The turbofan gas turbine engine 100 comprises in axial flow sequence, a heat exchanger module 110, a fan assembly 130, a compressor module 140, and a turbine module 150.

(14) In the present arrangement, the fan assembly 130 comprises two fan stages 131, with each fan stage 131 comprising a plurality of fan blades 132. In the present arrangement each fan stage 131 has the same fan diameter 136, with the respective plurality of fan blades defining a fan diameter of 0.9 m. In an alternative arrangement, the two fan stages 131 may have different fan diameters 136 each defined by the corresponding plurality of fan blades 132. As previously mentioned, the fan diameter (D) 136 is defined by a circle circumscribed by the leading edges of the respective plurality of fan blades 132.

(15) The heat exchanger module 110 comprises a plurality of heat transfer elements 112. The heat exchanger module 110 is in fluid communication with the fan assembly 130 by an inlet duct 160. The heat exchange module 110 has an axial length 115 of 0.4 m, this being 0.4 times the fan diameter of 0.9 m.

(16) The inlet duct 160 extends between a downstream-most face of the heat transfer elements and an upstream-most face of the fan assembly. In the present arrangement, the inlet duct 160 is linear. However, in other arrangements the inlet duct 160 may be curved or convoluted.

(17) The inlet duct 160 has a fluid path length 164 of 3.6 m, this being 4.0 times the fan diameter of 0.9 m. The fluid path length 164 extends along a central axis 162 of the inlet duct 160.

(18) As outlined earlier, the heat exchanger module 110 has a flow area (A.sub.HEX) 118. The heat exchanger module flow area 118 is the cross-sectional area of the heat exchanger module 110 through which an air flow 104 passes before being ingested by the fan assembly 130. In the present arrangement, the heat exchanger module flow area 118 has an annular cross-section and corresponds directly to the shape of the air flow passing through the heat exchanger module 110.

(19) The fan assembly 130 has a corresponding flow area (A.sub.FAN) 138. The fan assembly flow area 138 is the cross-sectional area of the fan assembly 130 through which an air flow 104 passes before separating into a core engine flow and a bypass flow. The fan assembly flow area 138 has an annular shape since it corresponds to the annular area swept by the fan blades 132.

(20) In the present arrangement, the heat exchanger module flow area 118 is equal to the fan assembly flow area 138, and the corresponding ratio of A.sub.HEX/A.sub.FAN is equal to 1.0.

(21) The heat exchanger module 110 has a flow diameter (E) 116, which is the diameter of the air flow passing through the heat exchanger module 110. In the present arrangement, the heat exchanger module flow diameter 116 is equal to the fan diameter 136.

(22) The heat exchanger module 110 comprises a plurality of heat transfer elements 112 for the transfer of heat energy from a first fluid 190 contained within the heat transfer elements 112 to an airflow 104 passing over a surface 113 of the heat transfer elements 112 prior to entry of the airflow 104 into the fan assembly 130. In the present embodiment, the first fluid 190 is a mineral oil. In other arrangements, the first fluid 190 may be an alternative heat transfer fluid such as, for example, a water-based fluid, or the fuel used by the turbofan gas turbine engine.

(23) The heat transfer elements 112 have a conventional tube and fin construction and will not be described further. In an alternative arrangement, the heat transfer elements may have a different construction such as, for example, plate and shell.

(24) The turbofan gas turbine engine 100 further comprises an outer housing 170. The outer housing 170 fully encloses the sequential arrangement of the heat exchanger module 110, inlet duct 160, fan assembly 130, compressor module 140, and turbine module 150. The outer housing 170 defines a bypass duct 180 between the outer housing 170 and the core engine components (comprising inter alia the compressor module 140 and the turbine module 150). In the present arrangement, the bypass duct 180 has a generally axi-symmetrical annular cross-section extending over the core engine components. In other arrangements, the bypass duct 180 may have a non-symmetric annular cross-section or may not extend around a complete circumference of the core engine components.

(25) FIG. 4 shows a perspective view of the heat exchanger module 110 and fan assembly 130 of the turbofan gas turbine engine 100 according to the first embodiment. The heat exchanger module 110 comprises sixteen radially extending vanes 120 arranged in a circumferential array 122. Alternative embodiments may have more or fewer radially extending vanes 120. In the present arrangement, each of the vanes 120 is hollow and each vane comprises a single heat transfer element 112 positioned within the vane 120. The alternative embodiments may not have a heat transfer element 112 within each vane 120 or may have more than one heat transfer element 112 in any single vane 120.

(26) Each of the heat transfer elements 112 has a corresponding swept area, which is the area of the heat transfer element 112 that is contacted by the air flow 104. In the present arrangement, the total swept heat transfer element area (A.sub.HTE) is the sum of the swept area of each of the individual heat transfer elements 112.

(27) Each vane 120 is configured to allow the air flow 104 passing through the heat exchange module to pass through the hollow portion of the vane 120 and thence to flow over the respective heat transfer element 112. In this way heat energy is transferred from the first fluid 190 to the air flow 104.

(28) Fan to Heat Transfer Element Area Parameter

(29) A Fan to Heat Transfer Element Area parameter F.sub.EA is defined as a ratio of the total swept heat transfer element area (A.sub.HTE—defined in the preceding paragraph) to the fan assembly flow area (A.sub.FAN—defined earlier). For the present arrangement, the swept heat transfer element area is approximately 52 m.sup.2, while the fan assembly flow area is approximately 0.43 m.sup.2. This makes the Fan to Heat Transfer Element Area parameter approximately 121.

(30) Heat Energy Rejection Performance

(31) In use, the first fluid 190 enters the heat transfer elements 112 having a maximum temperature of 80° C. The heat transfer module 110, comprising all of the heat transfer elements 112, transfers approximately 325 kW of heat energy from the first fluid 190 passing through the heat transfer elements 112 to the air flow 104 passing through the heat exchanger module 110.

(32) In the present application, the first fluid 190 draws heat energy from, for example, mechanical systems such as, for example, the engine lubrication system, and electrical systems both on the turbofan engine and external to the turbofan engine. In other arrangements, the first fluid 190 may draw heat energy only from the turbofan engine, or alternatively only from systems external to the turbofan engine.

(33) Heat Exchanger Performance Parameter

(34) In use, the turbofan gas turbine engine 100 according to the first embodiment has a maximum dry thrust of 190 kN at a full-power engine condition. The term ‘dry thrust’ is understood to mean the engine's thrust performance without any supplementary thrust such as from an exhaust reheat system or similar. The engine's thrust performance is measured at standard sea-level static (SLS) atmospheric conditions (i.e. 15° C., 1013 mbar).

(35) With the turbofan engine 100 operating at a full-power condition referred to above, the fan assembly 130 will have a maximum rotational speed of approximately 9500 rpm.

(36) As outlined above, the heat exchanger module 110 transfers approximately 325 kW of heat energy from the first fluid 190 to the airflow 104 passing through the heat exchanger module 110.

(37) Consequently, a Heat Exchanger Performance parameter P.sub.EX is defined as a ratio of the heat energy rejection to the maximum dry thrust. For the turbofan gas turbine engine of the present arrangement, the ratio P.sub.EX is 1.7.

(38) Referring to FIG. 3, a turbofan gas turbine engine according to a second embodiment of the disclosure is designated generally by the reference numeral 200. Features of the turbofan gas turbine engine 200 which correspond to those of turbofan gas turbine engine 100 have been given corresponding reference numerals for ease of reference.

(39) The turbofan gas turbine engine 200 comprises in axial flow sequence, a heat exchanger module 210, a fan assembly 130, a compressor module 140, and a turbine module 150.

(40) The fan assembly 130, compressor module 140, and turbine module 150 correspond directly to the those of the first embodiment described above.

(41) The heat exchanger module 210 comprises a plurality of heat transfer elements 212 and is also in fluid communication with the fan assembly 130 by an inlet duct 260. As in the first embodiment, the inlet duct 260 extends between a downstream-most face of the heat transfer elements and an upstream-most face of the fan assembly.

(42) The inlet duct 260 has a fluid path length 264 along a central axis 162 of the inlet duct 260 of 2.4 m, this being 2.7 times the fan diameter of 0.9 m.

(43) The heat exchanger module 210 has a flow area (A.sub.HEX) 218. As in the first embodiment, the heat exchanger module flow area 118 is annular in cross-section. However, in this arrangement the heat transfer elements 212 do not extend completely across that cross-section of the heat exchange module 210 that is available for the flow 104. In other words, there is a radially proximal portion of the cross-section of the heat transfer module across which there are no heat transfer elements 212.

(44) The fan assembly 130 has a flow area (A.sub.FAN) 138 that, as described above, has an annular shape corresponding to the annular area swept by the fan blades 132.

(45) In the present arrangement, despite the heat exchanger module flow area 218 having different dimensions to the fan assembly flow area 138, the heat exchanger module flow area 218 is equal to the fan assembly flow area 138. As for the first embodiment, the corresponding ratio of A.sub.HEX/A.sub.FAN is equal to 1.0.

(46) The heat exchanger module 210 has a flow diameter 216. The heat exchanger module flow diameter 216 is greater than the fan diameter 136.

(47) The turbofan gas turbine engine 200 further comprises an outer housing 270. As with the first embodiment described above, the outer housing 170 fully encloses the sequential arrangement of the heat exchanger module 210, inlet duct 260, fan assembly 130, compressor module 140, and turbine module 150. The outer housing 270 also defines an annular bypass duct 180 between the outer housing 170 and the core engine components

(48) In use the turbofan gas turbine engine 200 functions in the same manner as described above in relation to the turbofan gas turbine engine 100 of the first embodiment.

(49) 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.

(50) 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.

(51) 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.

(52) 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.