Supersonic aircraft turbofan

Abstract

A turbofan engine has an engine core including in flow series a compressor, a combustor and a turbine. The engine further has a fan located upstream of the engine core, has a supersonic intake for slowing down incoming air to subsonic velocities at an inlet to the fan formed by the intake, has a bypass duct surrounding the engine core, wherein the fan generates a core airflow to the engine core and a bypass airflow through the bypass duct, and has a mixer for mixing an exhaust gas flow exiting the engine core and bypass airflow exiting bypass duct. The engine further has a thrust nozzle rearwards of the mixer for discharging mixed flows, the thrust nozzle having a variable area throat. The engine further has a controller controlling the thrust produced by the engine over a range of flight operations including on-the-ground subsonic take-off and subsequent off-the-ground subsonic climb.

Claims

1. A turbofan engine for providing propulsive thrust to a supersonic aircraft, the turbofan engine comprising: an engine core including, in flow series, a compressor, a combustor, and a turbine; a fan located upstream of the engine core; a supersonic intake configured to decrease air velocities of incoming air to subsonic velocities at an inlet to the fan formed by the intake; a bypass duct surrounding the engine core, the fan being configured to generate a core airflow to the engine core and to generate a cold bypass airflow through the bypass duct; a variable mixer configured to mix a hot exhaust gas flow exiting the engine core with the cold bypass airflow exiting the bypass duct, the mixer being configured to vary a cross-sectional area of a hot inlet of the mixer for the hot exhaust gas flow and a cross-sectional area of a cold inlet of the mixer for the cold bypass airflow to change a ratio of an amount of the hot exhaust gas flow relative to an amount of the cold bypass airflow that is mixed by the mixer; a thrust nozzle located rearwards of the mixer, the thrust nozzle being configured to discharge the exhaust gas flow and the bypass airflow mixed by the mixer, the thrust nozzle having a variable area throat defining a cross-sectional throat area, the thrust nozzle being configured to controllably vary the throat area; and a controller configured to: perform a ground subsonic take-off operation by controlling the thrust nozzle to reduce the throat area of the variable area throat, which increases the thrust of the turbofan engine, and perform an off-the-ground subsonic climb operation by controlling the thrust nozzle to increase the throat area of the variable area throat.throat, and controlling the mixer to increase the cross-sectional area of the hot inlet and decrease the cross-sectional area of the cold inlet, which increase the amount of hot exhaust gas flow in the ratio of the hot exhaust gas flow relative to the cold bypass airflow, such that a value of (AHOT.sub.after/ACOLD.sub.after)/(AHOT.sub.before/ACOLD.sub.before) is in a range from 1.1 to 2.5, wherein AHOT is the cross-sectional area of the hot inlet of the mixer, ACOLD is the cross-sectional area of the cold inlet of the mixer, and subscripts “before” and “after” denote respectively immediately before and immediately after the relative increase in cross-sectional area of the hot inlet.

2. The turbofan engine according to claim 1, wherein the intake has a fixed geometry.

3. The turbofan engine according to claim 1, wherein the controller is configured to perform the increase of the cross-sectional throat area of the variable area throat at a first predetermined point during the climb operation.

4. The turbofan engine according to claim 3, wherein the first predetermined point is any one of a predetermined aircraft speed, a predetermined fan speed, a predetermined aircraft altitude, a predetermined aircraft flap position, aircraft landing gear retraction, and a predetermined time.

5. The turbofan engine according to claim 1, wherein the controller is configured to perform the increase of the cross-sectional throat area of the variable area throat by from 10% to 70% of the cross-sectional throat area immediately before performing the increase.

6. The turbofan engine according to claim 1, wherein during the take-off operation or the climb operation, the controller is configured to reduce fuel flow to the combustor.

7. The turbofan engine according to claim 6, wherein the controller is configured to perform the reduction of the fuel flow to the combustor at a second predetermined point during the take-off operation or the climb operation.

8. The turbofan engine according to claim 7, wherein the second predetermined point is any one of a predetermined aircraft speed, a predetermined weight-on-wheels, a predetermined distance along runway, a predetermined aircraft speed, a predetermined fan speed, a predetermined aircraft altitude, a predetermined aircraft flap position, aircraft landing gear retraction, and a predetermined time.

9. The turbofan engine according to claim 1, wherein the controller is further configured, in the event of an engine failure or malfunction after the reduction of the cross-sectional throat area of the variable area throat, to increase the cross-sectional throat area of the variable area throat.

10. A supersonic aircraft having the turbofan engine according to claim 1.

11. A method of operating a supersonic aircraft having a turbofan engine which provides propulsive thrust to the supersonic aircraft over a range of flight operations including a transonic push operation during which the supersonic aircraft transitions from subsonic flight to supersonic flight, and a supersonic cruise operation, which has a relatively lower thrust than the transonic push operation, the turbofan engine including: an engine core including, in flow series, a compressor, a combustor, and a turbine; a fan located upstream of the engine core; a supersonic intake configured to decrease air velocities of incoming air to subsonic velocities at an inlet to the fan formed by the intake; a bypass duct surrounding the engine core, the fan being configured to generate a core airflow to the engine core and to generate a cold bypass airflow through the bypass duct; a variable mixer configured to mix a hot exhaust gas flow exiting the engine core with the cold bypass airflow exiting the bypass duct, the mixer being configured to vary a cross-sectional area of a hot inlet of the mixer for the hot exhaust gas flow and a cross-sectional area of a cold inlet of the mixer for the cold bypass airflow to change a ratio of an amount of the hot exhaust gas flow relative to an amount of the cold bypass airflow that is mixed by the mixer; and a thrust nozzle located rearwards of the mixer, the thrust nozzle being configured to discharge the exhaust gas flow and the bypass airflow mixed by the mixer, the thrust nozzle having a variable area throat defining a cross- sectional throat area, the thrust nozzle being configured to controllably vary the throat area, the method comprising: performing a ground subsonic take-off operation by controlling the thrust nozzle to reduce the throat area of the variable area throat, which increases the thrust of the turbofan engine; and transitioning the supersonic aircraft after the take-off operation to perform an off-the-ground subsonic climb operation by controlling the thrust nozzle to increase the throat area of the variable area throat, and controlling the mixer to increase the cross-sectional area of the hot inlet and decrease the cross-sectional area of the cold inlet, which increase the amount of hot exhaust gas flow in the ratio of the hot exhaust gas flow relative to the cold bypass airflow, such that a value of (AHOT.sub.after/ACOLD.sub.after)/(AHOT.sub.before/ACOLD.sub.before) is in a range from 1.1 to 2.5, wherein AHOT is the cross-sectional area of the hot inlet of the mixer, ACOLD is the cross-sectional area of the cold inlet of the mixer, and subscripts “before” and “after” denote respectively immediately before and immediately after the relative increase in cross-sectional area of the hot inlet.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) Embodiments of the present disclosure will now be described by way of example with reference to the accompanying drawings in which:

(2) FIG. 1 shows a longitudinal cross-section through a ducted fan gas turbine main engine for supersonic flight operation of an aircraft; and

(3) FIG. 2A shows, for three aircraft flight speed design Mach numbers of M1.2, M1.8 and M2.2, plots of maximum take-off (MTO) thrust against area variabilities for thrust nozzle throat area (A8), the hot stream mixer area (Hot Area) and cold stream mixer area (Cold Area) relative to a fixed geometry engine at an exhaust jet velocity at a noise limit of 1100 ft/s (335 m/s) and a fixed turbine inlet temperature;

(4) FIG. 2B shows corresponding plots of the change in the ratio of cold stream velocity to hot stream velocity against area variabilities for the hot stream mixer area and the cold stream mixer area;

(5) FIG. 3 shows a contour plot illustrating a relationship between fan diameter, top of climb (TOC) thrust and fan pressure ratio (FPR), the plot being overlaid with contour lines (running downhill from left to right) representing achievable take-off thrust when the engine is run in a maximum thrust configuration, and with further contour lines (running much more steeply downhill from right to left) depicting achievable take-off thrust in a maximum allowed noise configuration;

(6) FIG. 4A shows thrusts at take-off (F_TO), sideline (F_SL) and fly-over (F_FO) noise points together with the take-off field length (x_lo) and fan diameter (DFAN) of a baseline engine (BASE), a first engine variant (PLR Area) and a second engine variant (PLR Dfan, Area, T41);

(7) FIG. 4B shows achievable flight trajectories for the three engines; and

(8) FIG. 4C shows estimations of the relative certification noise produced by the three engines, dSPL_SL (Vjet) being sideline noise, dSPL_CB (FN) being flyover noise and dSPL_cum being cumulative noise.

DETAILED DESCRIPTION OF THE DISCLOSURE

(9) 1. Engine

(10) FIG. 1 shows a longitudinal cross-section through a turbofan engine 1 for a civil supersonic aircraft. The turbofan engine 1 comprises an engine intake 2, a single stage fan 3 (although in other embodiments a multi-stage fan can be used), a primary flow channel 6 that leads through a engine core, a secondary flow channel 5 that leads past the engine core, a multi-lobed mixer 12, and a thrust nozzle 4 into which a thrust reverser 14 (shown in both stowed and deployed configurations) can be integrated.

(11) The turbofan engine 1 has a machine axis or engine centre line 8. The machine axis 8 defines an axial direction of the turbofan engine. A radial direction of the turbofan engine extends perpendicularly to the axial direction.

(12) The engine core comprises in a per se known manner a compressor 71, 72, a combustion chamber 11 and a turbine 91, 92. In the shown exemplary embodiment, the compressor comprises a booster compressor 71 and a high-pressure compressor 72. The turbine that is arranged behind the combustion chamber 11 comprises a high-pressure turbine 91 and a low-pressure turbine 92. The high-pressure turbine 91 drives a high-pressure shaft 81 that connects the high-pressure turbine 91 to the high-pressure compressor 7. The low-pressure turbine 92 drives a low-pressure shaft 82 that connects the low-pressure turbine 92 to the booster compressor 71 and the single stage fan 3.

(13) The turbofan engine 1 is arranged inside an engine nacelle 10. It is connected to the aircraft fuselage, for example via a pylon.

(14) The engine intake 2 forms a supersonic air inlet and is correspondingly provided and suitable for slowing down the inflowing air to velocities of below Ma 1.0. In FIG. 1, but not necessarily, the engine intake is bevelled, forming an angle α, wherein the upper edge projects with respect to the lower edge. This serves for a better distribution of compression shocks as they occur in supersonic flight. However, in principle the engine intake can be formed in a straight manner, i.e. with an angle α of 90°, or a different angle than the one shown.

(15) Upstream, the fan 3 is provided with a nose cone 35. Behind the fan 3, the flow channel through the fan 3 is divided into the primary flow channel 6 and the secondary flow channel 5. The secondary flow channel 5 is also referred to as the bypass flow channel or the bypass channel. A fan stator 32 comprising a plurality of stator blades is located in the secondary flow channel 5 behind the fan 3.

(16) Behind the engine core, the primary flow from the primary flow channel 6 and the secondary flow from the secondary flow channel 5 are mixed by the mixer 12. Further, an outlet cone 13 is inserted behind the turbine to realize the desired cross section of the primary flow channel.

(17) The thrust nozzle 4 has a variable area throat, e.g. as described in US 2004/0006969 A1 and U.S. Pat. No. 8,453,458. Further, the mixer 12 can be a variable area mixer which allows the relative areas available for the primary flow and the bypass flow at the mixer to be adjusted, and/or the outlet cone 13 can be adjustable to vary the area available for the relative areas available for the hot exhaust gas flow and the cold bypass airflow flow exiting the engine core at the mixer, e.g. as also described in US 2004/0006969 A1 and U.S. Pat. No. 8,453,458.

(18) Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have a different number of interconnecting shafts (e.g. one or three) and/or a different number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.

(19) 2. General Control Principles

(20) The engine has an EEC (not shown) which controls many aspects of the engine's performance. In particular, the EEC controls the engine cycle to decouple take-off thrust requirements on the ground with the following early stages of flight (i.e. initial climb).

(21) Specifically, a high exhaust jet velocity is produced to achieve the required take-off speed in a sensible take-off field length. The EEC obtains the elevated thrust in this phase of the mission (recognised by the EEC, for example, by an aircraft speed below take-off speed combined with aircraft on ground indicated by, for example, weight on the wheels) by reducing the thrust nozzle throat area while supplying a high fuel flow to the combustor (although remaining within the ultimate temperature limits of the engine). Reducing the thrust nozzle throat area while increasing fan speed so as to maintain or increase mass flow, increases the pressure ratio of the fan and thus increases the exhaust jet velocity.

(22) Upon leaving the ground (as indicated, for example, by no weight on wheels and/or aircraft speed in excess of critical take off velocity), the EEC reduces the jet velocity and hence the thrust to a level below that on the ground but adequate to achieve a minimum climb gradient as set by aircraft certification requirements (these are typically defined on the one engine inoperative (OEI) case and the number of engines). The reduction in exhaust jet velocity is achieved by increasing the thrust nozzle throat area. This reduces the fan pressure ratio and reduces the exhaust jet velocity.

(23) Optionally, at a pre-defined marker (which can be air speed, or zero weight on wheels) the EEC initiates a reduction in fuel flow to the combustor. This also decreases the thrust and thus reduces the acceleration of the aircraft, but not to the point where the aircraft no longer accelerates. The thrust response of the engine to changes in fuel flow/throttle position may be different to the thrust response to changes in nozzle throat area. Accordingly, it may be preferable to initiate the fuel flow change at a different time to the nozzle throat area change, to improve engine operability.

(24) A further option is for the EEC to actuate a variable area mixer 12 and/or an adjustable outlet cone 13 to alter the relative areas available for the hot exhaust gas flow and the cold bypass airflow flow exiting the engine core at the mixer in order to achieve a more equal ratio of hot flow to cold flow velocities so that the noise generate by the exhaust is reduced. In particular increasing the area for the hot flow relative to that for the cold flow tends to equalise the velocities.

(25) By implementing the above adjustments via the thrust schedule of the EEC, no pilot intervention is required. The EEC can also implement a failure protocol in which it reverts to maximum thrust by increasing fuel flow to the combustor and reducing the thrust nozzle throat area.

(26) 3. Detailed Analysis

(27) Take-off thrust F.sub.TO is set by the relationship between aircraft weight (MTOW), lift coefficient (C.sub.L) air density (ρ) wing area (S.sub.ref), rolling resistance coefficient (μ.sub.r), and lift to drag ratio (L/D) and available take off distance (x.sub.io).

(28) $F_{\mathrm{TO}}=\frac{1}{2}.Math.g.Math.\mathrm{MTOW}.Math.\left(\frac{2.Math.\mathrm{MTOW}}{C_L.Math.\rho .Math.S_{\mathrm{ref}}.Math.x_{\mathrm{lo}}}+\mathrm{MTOW}\left(\frac{1}{2.Math.L/D}+\frac{3.Math.{\mu }_r}{2}\right)\right)$

(29) For an aircraft of wing area S.sub.ref=200 m.sup.2, MTOW=80000 kg, L/D=8.5, C.sub.L=0.5 and assuming a rolling resistance μ.sub.r=0.02, then to achieve a 8000 ft (2400 m) take-off field length requires a take-off thrust of F.sub.TO≈21 klbf (93 kN)

(30) A fan sized to achieve this thrust with a fan face Mach number of M=0.65 and a hub-to-tip ratio of HTR=0.3 while retaining an exhaust jet velocity below ˜1100 ft/s (335 m/s) in order to achieve an acceptable noise would have a diameter of 65 inches (1650 mm).

(31) The thrust requirement in a second segment of climb (in accordance with climb requirements set out in Title 14, Chapter I, Subchapter C, § 25.111 of the Aeronautics and Space, Airworthiness Standards: Transport Category Airplanes, Federal Aviation Administration, US Department of Transportation as of 16 Jul. 2018—see https://www.ecfr.gov/cgi-bin/text-idx?node=14:1.0.1.3.11#se14.1.25_1111) is set by a need to meet a prescribed climb gradient in OEI failure cases. The thrust required may be expressed in terms of the gravitational constant (g) aircraft weight (MTOW), climb gradient (Y.sub.OEI), lift to drag in one engine inoperative condition (L/D.sub.OEI) and the number of engines (N.sub.eng).

(32) $F_{2\mathrm{ndSeg}}=F_{\mathrm{OEI}}=\frac{g.Math.\mathrm{MTOW}.Math.\left({\gamma }_{\mathrm{OEI}}.Math.\frac{L}{D_{\mathrm{OEI}}}+1\right)}{\frac{L}{D_{\mathrm{OEI}}}\left(N_{\mathrm{eng}}-1\right)}$

(33) Assuming a reduction in L/D for OEI conditions to L/D.sub.OEI=7.5 and a three engine configuration, the thrust requirement to retain level flight (Y.sub.OEI=0) in OEI conditions reduces to F.sub.2ndSeg˜11.7 klbf (52 kN). A fan sized by the same principles as above to this lower thrust requirement would have a diameter of only about 49 inches (1240 mm).

(34) In order to reduce noise over a flyover microphone it is permissible to reduce the engine thrust to a lower level retaining a minimum climb gradient with all engines operative. The thrust level is set by:

(35) $F_{\mathrm{Final}}=\frac{g.Math.\mathrm{MTOW}.Math.\left({\gamma }_{\mathrm{Fin}}.Math.\frac{L}{D}+1\right)}{\frac{L}{D}{N}_{\mathrm{eng}}}$

(36) In this phase of flight, while retaining γ.sub.Fin=1.5%, the thrust requirement reduces further to F.sub.Final˜7.8 klbf (35 kN).

(37) Changes to engine thrust are conventionally achieved by changes to the fuel flow to the combustor. However, changing the geometry of the engine also allows changes to the engine operating point to be achieved. The gas flow through an engine is primarily controlled by the area at the inlet to the turbine, and in the case of a turbofan engine also by the throat area of the final thrust nozzle. Variation of the turbine inlet area is mechanically challenging to achieve. However, by reducing the throat area of the final thrust nozzle while increasing the fan speed to retain the same mass flow, the pressure rise through the fan may be increased together with the jet velocity and therefore thrust. Higher pressure ratios at a fixed flow move the fan operating point closer to the stability limit of the fan. But by reducing the area for the hot exhaust gas flow (which generally also entails increasing the area for the cold bypass airflow) at the mixer the operating point of the fan may be moved away from the stability limit.

(38) FIG. 2A shows, for three aircraft flight speed design Mach numbers of M1.2, M1.8 and M2.2, plots of maximum take-off (MTO) thrust against area variabilities for the thrust nozzle throat area (A8), the hot stream mixer area (Hot Area) and the cold stream mixer area (Cold Area) relative to a fixed geometry engine at an exhaust jet velocity at the noise limit of 1100 ft/s (335 m/s) and a fixed turbine inlet temperature. FIG. 2B shows corresponding plots of the change in the ratio of cold stream velocity to hot stream velocity against area variabilities for the hot stream mixer area and the cold stream mixer area (variation of A8 had no significant impact on this ratio). The fixed jet velocity constraint prevents an increase in fan pressure ratio and as a result increasing A8 results in a significant increase in thrust relative to the datum nozzle area as more mass flow passes through the fan. For the M2.2 engine more than 120% of the zero variability MTO thrust is achieved with a thrust nozzle throat area of 115% of the nominal value (the area designed for maximum engine air-flow flight condition such as at top of climb). Additionally varying the mixer hot and cold areas changes the relative velocities from the hot and cold streams allowing equalisation of these velocities to reduce the noise generated. Increasing the hot mixer area tends to equalise the stream velocities. Decreasing the cold mixer area also tends to equalise the stream velocities. FIG. 3 then shows a contour plot illustrating the relationship between fan diameter, top of climb (TOC) thrust and fan pressure ratio (FPR). Overlaid are contour lines (running downhill from left to right) representing the achievable take-off thrust when the engine is run in a maximum thrust configuration to a limiting turbine inlet temperature T41=T41max with the thrust nozzle throat area A8 at 107.5% of the design area and the hot mixer area AHOT at 100% of the design area. For these relationships, T41max is the same as the maximum rotor inlet temperature (RIT) and is taken as 1750 K. The design area is set by the maximum engine air-flow condition which occurs at top of climb conditions. Further contour lines (running much more steeply downhill from right to left) depict achievable take-off thrust in a maximum allowed noise configuration in which there is a noise limit constraint on exhaust jet velocity of 1100 ft/s (335 m/s). In this configuration A8 is 115% of the design value and AHOT is 130% of the design area. It is apparent that an engine of 50 inch (1270 mm) diameter delivering a top of climb thrust of 7000 lbf (31 kN) (indicated on FIG. 3 by the arrowed circle) is capable of achieving about 13750 lbf (161 kN) of thrust in the maximum allowed noise configuration or about 22500 lbf (100 kN) of thrust in the maximum thrust configuration.

(39) The thrust augmentation achievable on the ground relative to in-flight may be used to reduce the fan diameter at an approximately equal noise level for certification. FIG. 4A shows thrusts at take-off (F_TO), sideline (F_SL) and fly-over (F_FO) noise points together with the take-off field length (x_lo) and the fan diameter (DFAN) of a baseline (BASE) and two further engine variants. The first (PLR Area) uses only area variation of the thrust nozzle throat to achieve a 120% thrust augmentation. The second (PLR Dfan, Area, T41) employs throttle increase and thrust nozzle throat area variation to provide a 165% increase in thrust on the ground. FIG. 4B then shows achievable flight trajectories for the three engines, and FIG. 4C shows estimations of the relative certification noise produced by engines, dSPL_SL (Vjet) being sideline noise, dSPL_CB (FN) being flyover noise and dSPL_cum being cumulative noise. The area variation alone (PLR Area) delivers an engine of reduced diameter but by the certification noise measure lacks sufficient thrust to achieve sufficient height over the noise measuring positions, thereby resulting in an overall increase in certification noise. The variant including area variation and increased fuel flow achieves a cumulative noise margin equivalent to that the baseline, by trading sideline and flyover noise levels, but has a significantly reduced fan diameter allowing significant increase in aircraft range.

(40) Taking a nominal value for the thrust nozzle throat area as 1.0, then the reduced throat area at take-off typically has a value of about 0.9 for all types of aircraft (i.e. design Mach numbers), while the increased area of the variable area throat during subsequent climb varies between aircraft types, e.g. from about 1.12 to about 1.2. Thus, across aircraft types, based on these values the increase in the variable area throat typically lies in a range of from 10% to 70% of its area (i.e. 0.9) immediately before the increase.

(41) Similarly, it is possible to determine a typical range for the ratio (AHOT.sub.after/ACOLD.sub.after)/(AHOT.sub.before/ACOLD.sub.before), where AHOT is the area for the hot exhaust gas flow at the mixer, ACOLD is the area for the cold bypass airflow at the mixer, and the subscripts “before” and “after” denote respectively immediately before and immediately after a relative increase in the area for the hot exhaust gas flow relative to the area for the cold bypass airflow. In particular, if AHOT.sub.before and ACOLD.sub.before have nominal values of 1.0, then: If just AHOT.sub.after is varied, a maximum value for AHOT.sub.after is about 1.5, such that the ratio=1.5, and a minimum value for AHOT.sub.after is about 1.1, such that the ratio=1.1; and If just ACOLD.sub.after is varied, a minimum value for ACOLD.sub.after is about 0.6, such that the ratio=1.66, and a maximum value for ACOLD.sub.after is about 0.9, such that the ratio=1.11

(42) An extreme variation of a maximum value for AHOT.sub.after and a minimum value for ACOLD.sub.after thus gives a ratio of 1.5/0.6=2.5. Thus based on these values the ratio typically lies in a range of from 1.1 to 2.5, with a typical desired target ratio of 1.25/0.88=1.42 being based on AHOT.sub.after=1.25 and ACOLD.sub.after=0.88.

(43) 4. Summary

(44) By varying the thrust nozzle throat area after take-off and during the climb it is possible to achieve: A high take-off speed with high thrust within a reasonable field length A lower fan diameter while meeting noise constraints as a result of reduced thrust after take-off, the lower fan diameter providing a significant reduction in engine cross-sectional area and hence a significant reduction in wave drag on the aircraft and an increased in range

(45) While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Moreover, in determining extent of protection, due account shall be taken of any element which is equivalent to an element specified in the claims. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention.