Aircraft engine takeoff and climb thrust control

Abstract

The controller controls the aircraft's engines during the mission, which includes the takeoff phase, climb phase, and cruise phase, where the climb phase includes the peak climb (ToC) period before transitioning to the cruise phase. The controller can control one or more engines to output takeoff thrust during the takeoff phase, corresponding to the maximum thrust during the takeoff phase. The controller can further control one or more engines to output peak climb thrust during the ToC period, corresponding to the average thrust during the ToC period, wherein one or more engines are controlled such that the thrust ratio is between 1.1 and 1.6, defined as the ratio of the ToC thrust normalized to ToC environmental static pressure to the takeoff thrust normalized to takeoff environmental static pressure.

Description

Technical Field

[0001] This disclosure generally pertains to the operation of aircraft engines. Background Technology

[0002] In aviation, the aspect ratio of an aircraft wing refers to the ratio of its wingspan to its mean chord length. It is equal to the square of the wingspan divided by the wing area. The wing's aspect ratio, along with other platform characteristics, affects its aerodynamic efficiency because the lift-to-drag ratio increases with the aspect ratio, thus improving fuel economy. Traditionally, the use of high aspect ratio wing aircraft has been limited to small propeller-driven aircraft cruising at relatively low altitudes and speeds. Making high aspect ratio wing aircraft suitable for commercial flight still faces numerous design and operational challenges. Attached Figure Description

[0003] The complete and feasible disclosure of this disclosure, including its best mode, is set forth in the specification with reference to the accompanying drawings, for those skilled in the art, wherein:

[0004] Figure 1 Including diagrams of aircraft according to some embodiments;

[0005] Figure 2 Includes a flowchart of a method for controlling the thrust output of an aircraft engine according to some embodiments;

[0006] Figure 3 Including illustrations of truss-supported wings according to some embodiments;

[0007] Figure 4 Including a top view of a high aspect ratio wing according to some embodiments; and

[0008] Figure 5 Including schematic diagrams of a pipeless turbine engine according to some embodiments.

[0009] The elements in the accompanying drawings are for illustrative purposes only and are not necessarily drawn to scale. For example, the dimensions and / or relative positions of some elements in the drawings may be enlarged relative to other elements to facilitate a better understanding of the various embodiments of this teaching. Furthermore, to more clearly illustrate these different embodiments of this teaching, conventional elements that are practical or necessary in commercially viable embodiments but are well-known are generally not shown. Certain actions and / or steps may be described or illustrated in a particular order, and those skilled in the art will understand that a specific limitation regarding this order is not practically necessary. Detailed Implementation

[0010] The present disclosure will now be described in detail with reference to embodiments thereof, one or more of which are illustrated in the accompanying drawings. Each example is provided to explain the disclosure and not to limit it. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made to the disclosure without departing from its scope or spirit. For example, features shown or described as part of one embodiment may be used in combination with another embodiment to produce yet another embodiment. Therefore, the present disclosure is intended to cover such modifications and variations that fall within the scope of the appended claims and their equivalents.

[0011] The terms “first,” “second,” and “third” used in this article are interchangeable to distinguish one component from another, and are not intended to indicate the position or importance of the components.

[0012] Unless otherwise stated herein, the terms “connection,” “fixed,” “attached to,” and similar expressions refer to both direct connection, fixation, or attachment, and indirect connection, fixation, or attachment achieved through one or more intermediate components or features.

[0013] Unless the context clearly specifies otherwise, the singular forms “one,” “a,” and “this” include the plural referent.

[0014] The approximate terms used herein throughout the specification and claims are intended to modify any numerical representation that allows for variation without altering the essential function. Therefore, numerical values ​​modified by terms such as “about,” “approximately,” “almost,” and “substantially” are not limited to the specified precise values. In some cases, approximate terms may correspond to the precision of the instrument used to measure the value. For example, approximate terms may refer to an error range of 1%, 2%, 4%, 10%, 15%, or 20%. These approximate errors may apply to a single numerical value, to one or both endpoints of a range of values, and / or to the error in the range between endpoints. In this document and throughout the specification and claims, range limits may be combined and interchanged, and such ranges should be defined and include all subranges therein, unless otherwise indicated by context or text. For example, all ranges disclosed herein include endpoints, and the endpoints may be combined independently of each other.

[0015] Throughout this document and throughout the specification and claims, scope limits may be combined and interchanged, and unless otherwise indicated by context or wording, such scopes shall be deemed to include all subscopes thereof. For example, all scopes disclosed herein include endpoints, and these endpoints may be combined independently of each other.

[0016] The terms “front” and “rear” refer to relative positions within a gas turbine engine or aircraft, based on the normal operating attitude of the gas turbine engine or aircraft. More specifically, in this document, “front” and “rear” are used with reference to the direction of the aircraft’s flight and the direction of the gas turbine engine’s propulsion thrust. The terms “upstream” and “downstream” refer to relative directions of fluid flow in a fluid channel. For example, “upstream” refers to the direction of incoming fluid flow, and “downstream” refers to the direction of outgoing fluid flow.

[0017] The term “leading edge” refers to a component and / or surface that is generally oriented upstream relative to the fluid flow in the system; the term “trailing edge” refers to a component and / or surface that is generally oriented downstream relative to the fluid flow in the system.

[0018] "Blade" can refer to a stationary or rotating blade. "Stationary blade" and "wheel blade" have the same meaning.

[0019] The term "ductless aircraft engine" as used in this article refers to an aircraft engine characterized by having a rotating fan blade array and a stationary (non-rotating) exit guide vane (OGV) disposed behind the rotating fan blade array; or having a rotating fan blade array and a stationary, ductless inlet guide vane (IGV) disposed in front of the rotating fan blade array. In either case, the fan blades, inlet guide vane, or exit guide vane are not surrounded by a duct or fan nacelle.

[0020] During a flight mission, an aircraft typically goes through the takeoff phase, climb phase, cruise phase, descent phase, approach phase, and landing phase.

[0021] "Takeoff" (TO) refers to the flight phase in which an aircraft transitions from a stationary state on the ground to a state of flight in the air and ends when it reduces its throttle to enter the climb phase. The climb phase can occur at 10%–20% of the cruise altitude. During takeoff, the engines operate at maximum or near-maximum thrust in takeoff mode to generate the required speed and lift.

[0022] "Climb" refers to the phase of flight after takeoff where the aircraft ascends to a designated cruising altitude. During the climb phase, the engines continue to operate at higher thrust, but typically less than the maximum takeoff thrust. A climb phase can be characterized by a rate of climb exceeding 500, 1000, or 1500 feet per minute (fpm). In some embodiments, the rate of climb during the climb phase can be between 2500 and 3500 feet per minute.

[0023] "Cruise" or "cruise flight" refers to the flight phase in which an aircraft maintains a constant altitude ("cruise altitude") after the climb phase and before the descent phase. During the cruise phase, the aircraft engines operate in cruise mode, typically at 50% to 90% of their rated speed, for example, 70% to 80% of their rated speed. In various flight envelopes, cruise operation mode is defined as the operating mode of the gas turbine engine at a specific midpoint of the flight envelope based on the total fuel consumption of that flight envelope (e.g., when the gas turbine engine has consumed 50% of its total fuel during flight operation). During cruise, operating conditions can be between 50% and 100% of the engine's maximum effective altitude, where the aircraft's rate of climb or descent is less than + / - 500 feet per minute. Maximum effective altitude generally refers to the highest altitude at which the aircraft can maintain level flight under a given power setting and load, and is defined by the aircraft manufacturer in the aircraft specifications.

[0024] In several examples, cruise flight can be conducted at a maximum cruising altitude of approximately 65,000 feet (ft.). In some examples, the cruising altitude ranges from approximately 28,000 feet to approximately 45,000 feet. In other examples, the cruising altitude is expressed in flight levels (FL) based on sea-level standard pressure, with cruise flight occurring between FL280 and FL650. In yet another example, cruise flight occurs between FL280 and FL450. In some examples, the cruising altitude is defined at least based on atmospheric pressure, with cruise altitudes ranging from approximately 4.85 pounds per square inch (psia) to approximately 0.82 psia, based on a sea-level pressure of approximately 14.70 psia and a sea-level temperature of approximately 59 degrees Fahrenheit. In yet another example, the cruising altitude range ranges from approximately 4.85 psia to approximately 2.14 psia. It should be understood that in some examples, the pressure-defined range of cruising altitude may be adjusted based on different reference sea-level pressures and / or sea-level temperatures.

[0025] "Crest of Climb" (ToC) refers to the period during the climb phase of a flight just before entering the cruise phase. During ToC, the aircraft's rate of climb transitions from its peak rate of climb to the lower rate of climb required for the cruise phase. In one example, ToC is the last thousand or two thousand feet of altitude gain before reaching cruise altitude. In another example, ToC is the last two minutes of the climb phase before transitioning to the cruise phase.

[0026] "ToC thrust" refers to the average thrust generated by the engine during the ToC period.

[0027] "ToC (Total Cryop) environmental static pressure" refers to the average environmental static pressure measured during the ToC period. For example, the summit of the climb may occur at an altitude of 30,000 to 40,000 feet, at which point the ToC environmental static pressure may range from 4.373 psia to 2.730 psia.

[0028] "TO thrust" refers to the maximum thrust generated by the engine during the takeoff phase of flight. This thrust is usually equal to or close to the maximum thrust listed on the engine type certificate data sheet, and greater than the average thrust during the cruise phase.

[0029] "TO (Atmospheric Pressure)" refers to the atmospheric pressure measured at the aircraft's takeoff position during the takeoff phase of flight, typically between 11 and 15 psi.

[0030] The aspect ratio of an airfoil is the ratio of its span to its average chord width, which can be calculated by dividing the square of the wingspan by the projected area of ​​the airfoil.

[0031] High aspect ratio (HARW) and ultra-high aspect ratio (UHARW) aircraft are active areas of development for next-generation transport aircraft with increased fuel efficiency and performance. Generally, HARW refers to aircraft with an aspect ratio between 10 and 20, while UHARW refers to aircraft with an aspect ratio greater than 20. HARWs and UHARWs can employ strut-supported or truss-supported structures to support the wingspan and control bending moments. Because such designs reduce drag and increase lift in HARW and UHARW designs, these aircraft can take off at lower speeds and climb with less component stress and greater efficiency.

[0032] This disclosure generally describes aircraft engine control parameters suitable for HARW aircraft, which fully utilize the advantages of the HARW design by controlling engine thrust output at a thrust ratio that improves fuel efficiency during different flight phases. Although this document is primarily described with respect to HARW, in some embodiments, this disclosure is also applicable to the engines of UHARW aircraft.

[0033] In several aspects, this disclosure provides a method for controlling the operation of one or more engines of an aircraft during a flight mission, such that the thrust ratio between the ToC thrust normalized to the ToC environmental static pressure and the TO thrust normalized to the TO environmental static pressure generated by each of the one or more engines is in the range of 1.1 to 1.6. In other examples, this range is 1.13 to 1.5. The TO thrust and ToC thrust of the engine are partly limited by the design and size of the engine components. For example, the TO thrust is typically equal to or close to the engine's maximum rated thrust output. Therefore, in some embodiments, the core engine and fan of the engine used in the HARW aircraft are sized based on the thrust ratio described herein. In some embodiments, the thrust ratio can be achieved by a flight controller, such as an autothrottle system, and / or an engine controller, such as a full authority digital engine controller (FADEC).

[0034] As will be described herein, operating the engine according to the aforementioned thrust ratio means that, compared to conventional commercial aircraft engines, the engine's ToC thrust is closer to its TO thrust and maximum rated thrust. An engine designed and operated with this thrust ratio enables the HARW aircraft to climb at higher speeds and with greater fuel efficiency. This operating method allows the aircraft to reach cruise altitude in a shorter time, thereby extending the duration of the more fuel-efficient cruise phase.

[0035] The present disclosure will now be described in detail with reference to embodiments thereof, and one or more examples of the present disclosure are illustrated in the accompanying drawings. The specific description uses numbers and letters to refer to features in the drawings. The same or similar reference numerals in the drawings and specification are used to refer to the same or similar parts of the present disclosure.

[0036] For reference Figure 1 An example of an aircraft 120 is described below. The aircraft includes a fuselage 122, a pair of wings 124a and 124b extending from the fuselage 122, a flight controller 128, an engine controller 130 communicatively connected to the flight controller 128, and engines 126a and 126b structurally coupled to the wings 124a and 124b (engines 126a and 126b are further communicatively connected to the engine controller 130). Alternatively or additionally, engines 126a and 126b may also be coupled to the fuselage 122.

[0037] The fuselage 122 is the main body of the aircraft 120, used to house crew, passengers, cargo, or other items. The fuselage 122 is connected to the stabilization and lift elements of the aircraft 120, such as wings 124a and 124b. In some embodiments, the aircraft has a passenger capacity of 270 or fewer. That is, the fuselage is sized and arranged to accommodate 270 or fewer passengers and may be a narrow-body and / or single-aisle passenger aircraft.

[0038] Wings 124a and 124b are attached to fuselage 122 and provide lift to aircraft 120. The wings of the aircraft can be HARW or UHARW, with an aspect ratio greater than 10, for example, between 10 and 20.

[0039] In some embodiments, a truss connects each wing 124a, 124b to the fuselage 122. This truss stabilizes the wings and allows the aircraft to have larger engines coupled to the wings than typically would. A schematic diagram of the truss supporting the wing is shown below. Figure 3 As shown. In some embodiments, the wings 124a, 124b of the aircraft may include folding wingtips that can move up or down or fold via hinges. For example, the wingtips of the HARW aircraft can be folded when entering a terminal building, allowing the HARW aircraft to operate within spaces designed for conventional aircraft with smaller wingspans.

[0040] Engines 126a and 126b can be either ductless or ducted aircraft engines. Aircraft 120 typically includes two or more engines, such as two, three, or four. The specific configuration described herein is particularly suitable for ductless aircraft engines. Turbofan engines operate on the principle of a central gas turbine core driving a bypass fan located radially between the fan duct and the engine core. Ductless propulsion systems operate on the principle that the bypass fan is not located inside the fan duct. Compared to a bypass fan located inside a fan duct, removing the fan duct allows for the use of larger fan blades, enabling the application of a larger volume of air. Ductless propulsion systems can achieve higher propulsion efficiency compared to ducted turbofan engines. This document will combine... Figure 5 Examples of ductless turbofans are described. The engine may be an open rotor engine, comprising a rotor and stator blades located behind the rotor. The engine may be an open rotor engine, comprising counter-rotating rotors. The engine may further include an annular core inlet. In some embodiments, the engine may be a counter-flow engine.

[0041] In some embodiments, each of one or more engines 126a, 126b produces a maximum thrust of less than 50,000 pounds (approximately 222,400 Newtons). In some embodiments, each of one or more engines 126a, 126b is configured to have a bypass ratio (BPR) of 15 to 100. In some embodiments, the engine's BPR is 10.0 or higher. The BPR of a turbofan engine is the ratio of the bypass mass flow rate (flowing around the engine core) to the mass flow rate entering the engine core. For example, a BPR of 10:1 means that for every 1 kg of air flowing through the core, 10 kg of air flows through the bypass duct. In some embodiments, engines 126a, 126b are configured to have a fan pressure ratio (FPR) of 1.05 to 1.5. The FPR is the ratio of the pressure of air after it has passed through the fan to the pressure of air before it enters the fan. In a turbofan engine, the FPR is a variable that affects a given fuel consumption rate (SFC). For a given BPR, there exists an optimal FPR that minimizes the SFC. That is, FPR is usually chosen to balance BPR, thereby improving fuel efficiency.

[0042] Aircraft 120 includes a flight controller 128 that provides instructions to engine controller 130 during flight missions. Flight controller 128 may be any microcontroller, computer, or processor-based device (or multiple such elements) having a processor, memory, and programmable input / output peripherals, typically designed to control the operation of other components and devices. In some embodiments, aircraft 120 includes an autopilot system that automatically performs control during one or more phases of flight, including takeoff, climb, cruise, descent, approach, and / or landing. In some embodiments, flight controller 128 also includes an autothrottle system that receives instructions from the autopilot system and / or pilot interface to provide target engine thrust to engine controller 130 to maintain or reach a predetermined altitude and speed. In some embodiments, the autopilot system and / or autothrottle system may also provide mode selection to engine controller 130 to indicate the current flight phase (e.g., takeoff, climb, cruise, descent, landing, etc.). In some embodiments, the flight controller 128 includes a pilot interface located in the cockpit to allow the crew of the aircraft 120 (e.g., the pilot and / or co-pilot) to control the flight operations of the aircraft 120, and the autothrottle system can provide instructions to the engine controller 130 based on the control of the pilot interface.

[0043] Engine controller 130 may be any microcontroller, computer, or processor-based device (or multiple such elements) having a processor, memory, and programmable input / output peripherals, and is generally designed to control the operation of other components and devices. In some embodiments, engine controller 130 includes a full-authority digital engine controller (FADEC). In some embodiments, a separate FADEC controller may be assigned to each engine 126a, 126b. Engine controller 130 is generally configured to perform the methods described herein, including, when the methods are embodied in software program form, executing... Figure 2 The method shown. That is, an instruction may be stored on a computer-readable storage medium, which, when executed by a processor of the aircraft and / or engine controller 130, causes the processor to perform a reference... Figure 2 The method described above for operating one or more engines in an aircraft.

[0044] Engine controller 130 is configured to receive instructions from flight controller 128 and adjust the operation of engines 126a, 126b to conform to or execute those instructions. When flight controller 128 issues a target thrust command, engine controller 130 converts these commands into electrical commands to alter the operation of one or more thrust actuators 131 of engines 126a, 126b, thereby generating thrust. Thrust actuator 131 generally refers to an engine component capable of altering the engine's thrust output. In some embodiments, thrust actuator 131 may include one or more pitch fan blades, a fuel flow controller, multiple variable stator blades, multiple inlet guide vanes, multiple outlet guide vanes, variable nozzles, an electric motor, etc. Details of engine components that may function as thrust actuators will be referenced in [reference]. Figure 5 A more detailed description is provided below. In some embodiments, the engine controller 130 may also determine control of one or more thrust actuators 131 based on mode selection received from the flight controller 128 to improve engine efficiency. For example, the selection of a particular thrust actuator 131 that is altered to achieve a target thrust may differ based on the mode selection.

[0045] In some embodiments, the flight controller 128 and / or engine controller 130 are configured to operate engines 126a, 126b such that the ratio of the ToC thrust (normalized to the ToC ambient static pressure associated with the climb peak) generated by each of the one or more engines 126a, 126b at the ToC is in the range of 1.1 to 1.6 to the takeoff (TO) thrust (normalized to the TO ambient static pressure associated with takeoff) generated by the one or more engines 126a, 126b at takeoff is in the range of 1.1 to 1.6. In other examples, this range is 1.13 to 1.5. In some examples, the relationship between ToC altitude and climb rate allows the time to reach the climb peak to be less than 40 minutes. The ToC period typically ends at an altitude above 28,000 feet. In some embodiments, the flight controller 128 and / or engine controller 130 may be physically separate processor-based devices or a single integrated control unit.

[0046] For reference Figure 2 This describes an example of a method for operating an aircraft engine. In some embodiments, Figure 2 The steps in this process can be performed by one or more processor-based devices to control one or more thrust actuators of one or more aircraft engines. In some embodiments, Figure 2 The steps can be performed by the aircraft 120 and / or engine controller 130 via direct or indirect control, such as... Figure 1 The one or more thrust actuators 131 described above perform these steps. These steps may be embodied in software and stored in the form of computer-executable code on a computer-readable storage medium for execution by a processor, such as a processor in the aircraft 120 and / or the engine controller 130.

[0047] In step 132, the aircraft is in the takeoff phase, and the engines operate in takeoff mode. During takeoff, the engines are controlled to generate TO thrust, which reaches or approaches the engine's maximum rated thrust to obtain the lift required for takeoff. For HARW aircraft, takeoff lift can be achieved at lower speeds compared to conventional commercial aircraft with shorter wingspans. Therefore, the TO thrust of HARW aircraft can also be lower than that of conventional commercial aircraft. In some embodiments, the TO thrust range can be 25,000 to 40,000 pounds for single-aisle passenger aircraft with a capacity of 100–330 passengers, and 8,000 to 26,000 pounds for regional jets with a capacity of 50–160 passengers.

[0048] In step 134, the aircraft lifts off and enters the climb phase to gain altitude. The engines also switch from takeoff mode to climb mode. During the climb phase, the engine thrust output may be gradually reduced to maintain the target climb rate. The climb phase includes an initial climb period and a ToC period preceding the cruise phase. The average thrust of the engine during the ToC period is called the ToC thrust. In some embodiments, the ToC thrust ranges from less than 90% of the takeoff thrust to greater than the cruise thrust.

[0049] In steps 132 and 134, the engine operates at a thrust ratio defined by the following formula:

[0050]

[0051] Wherein, Fn_ToC is the average thrust during the ToC period; Ps_ToC is the ambient static pressure during the ToC period; Fn_TO is the maximum thrust generated by the engine at takeoff, which can be the engine's maximum rated thrust; and Ps_TO is the ambient static pressure during takeoff. The two thrust values ​​(Fn_ToC and Fn_TO) are normalized using their corresponding ambient static pressures (Ps_ToC and Ps_TO) to eliminate the influence of cruise altitude changes in the ratio.

[0052] In steps 132 and 134, the engine is controlled to generate TO thrust and ToC thrust, with a thrust ratio between 1.2 and 1.6. In some embodiments, the thrust ratio is between 1.13 and 1.5.

[0053] In step 136, the aircraft completes the climb phase and enters the cruise phase. In some embodiments, the process can proceed from step 134 to step 136 when the flight controller detects a target cruise altitude (e.g., based on altimeter data) and issues a command to put the engines into cruise mode. During the cruise phase, the engines operate at a steady thrust to maintain the aircraft's cruise speed. The engine controller is typically configured to control one or more engines to enable the aircraft to cruise at speeds of Mach 0.5 and above.

[0054] After step 136, the flight can continue into the descent, approach, and landing phases. These subsequent phases do not affect the thrust ratio described herein and are therefore omitted for brevity.

[0055] Referring now to Figure 3, a schematic diagram of a wing with a truss support structure is shown. Combined with... Figure 1 and Figure 2 The system and method described herein can be applied to aircraft equipped with truss-supported wings.

[0056] In this example, engine 126a is mounted on wing 124a, which is attached to fuselage 122. Truss 152 extends from fuselage 122 to wing 124a to support and stabilize the wing during operation of aircraft 120. Truss 152 may be made of suitable materials such as metal or alloy. Stud 150 is attached from truss 152 to wing 124a to provide further support and stability. The use of truss 152 and stud 150 can also increase the size or weight of the engine mounted on wing 124a. Truss 152 is inclined upward relative to wing 124a at an angle θ. A channel 154 is formed between truss 152 and wing 124a, and engine 126a is arranged in this opening area. It should be understood that truss 152 can have various shapes, sizes, and structural forms. For example, wing 124a may be a low-mounted wing, in which case truss 152 may be connected to the upper part of fuselage 122.

[0057] For reference Figure 4 This describes an example of a top view of an aircraft wing that can employ multiple aspect ratios. For example... Figure 4 As shown, wing 124a has a wingspan b, a standard mean chord length (SMC), and a projected area S. The wing aspect ratio AR is the ratio of the square of the wingspan b to the projected area S, and it is also equal to the ratio of the wingspan b to the standard mean chord length. For HARW aircraft operating with the thrust ratio described herein, the aspect ratio can be greater than 10, or between 10 and 20. The technical solutions disclosed herein are also applicable to UHARW aircraft with an aspect ratio greater than 20.

[0058] In some examples, the wingtip 123 can fold upwards along the fold line 121. This upward-folding wing design has advantages, allowing the aircraft to operate in space-constrained airports. Due to the high aspect ratio of the wing, if it were not folded, its length would exceed the capacity of the narrow boarding areas at many airports.

[0059] Figure 5 A schematic cross-sectional view of a gas turbine engine is provided, which can be applied to any of the schemes described herein, or used as any engine described herein. Specifically, Figure 5 The engine shown includes a rotor assembly with a single-stage ductless rotor blade. Accordingly, this rotor assembly may be referred to herein as a "ductless fan," or the entire engine 600 may be referred to herein as a "ductless aircraft engine." Furthermore, Figure 5 The engine shown also features a third airflow that extends from the compressor section to the rotor assembly flow path above the turbomachinery, which will be described in more detail below.

[0060] For reference, engine 600 is defined by an axial direction A, a radial direction R, and a circumferential direction C. Furthermore, engine 600 is defined by an axial centerline or central axis 612 extending along the axial direction A. Typically, the axial direction A extends parallel to the central axis 612, the radial direction R extends toward and away from the central axis 612 in a direction orthogonal to the axial direction A, and the circumferential direction extends 360° around the central axis 612. Engine 600, for example, extends along the axial direction A between a front end 614 and a rear end 616.

[0061] Engine 600 includes a turbine 620 and an upstream rotor assembly (also referred to as fan section 650). Typically, turbine 620 includes, in series along the airflow direction, a compressor section, a combustion section, a turbine section, and an exhaust section, and defines an annular core inlet 624. Core engine cowling 622 also at least partially surrounds the low-pressure and high-pressure systems. For example, the illustrated core engine cowling 622 at least partially surrounds and supports a booster stage or low-pressure (LP) compressor 626 for pressurizing the air entering turbine 620 through an annular core inlet 624. A high-pressure (HP) multistage axial compressor 628 receives the pressurized air from the low-pressure compressor 626 and further increases the air pressure. The pressurized air flows downstream to the combustion chamber 630 of the combustion section, where fuel is injected into the pressurized air and ignited to increase the temperature and energy level of the pressurized air.

[0062] It should be understood that the terms "high / low speed" and "high / low pressure" used in this document are interchangeable for high-pressure / high-speed systems and low-pressure / low-speed systems. Furthermore, it should be understood that the terms "high" and "low" are used in the same context to distinguish between the two systems and do not refer to any absolute speed and / or pressure values.

[0063] High-energy combustion products flow downstream from combustion chamber 630 to high-pressure turbine 632. High-pressure turbine 632 drives high-pressure compressor 628 via high-pressure shaft 636. In this configuration, high-pressure turbine 632 is drivably connected to high-pressure compressor 628. Subsequently, high-energy combustion products flow to low-pressure turbine 634. Low-pressure turbine 634 drives low-pressure compressor 626 and related components of fan section 650 via low-pressure shaft 638. In this configuration, low-pressure turbine 634 is drivably connected to low-pressure compressor 626 and components of fan section 650. In this embodiment, low-pressure shaft 638 and high-pressure shaft 636 are coaxially arranged. After driving each turbine 632, 634, combustion products are discharged from turbine 620 through turbine exhaust nozzle 640.

[0064] Therefore, turbine 620 defines a working gas flow path, or core duct 642, which extends between annular core inlet 624 and turbine exhaust nozzle 640. Core duct 642 is an annular duct, located approximately inside the core turbine cowling 622 in the radial direction R. Core duct 642 (e.g., through the working gas flow path of turbine 620) may be referred to as a second airflow.

[0065] Fan section 650 includes fan 652, which in this embodiment is the main fan. For Figure 5 In the illustrated embodiment, fan 652 is an open rotor or ductless fan. As shown, fan 652 includes a set of fan blades 654 (only one is shown in the figure). The fan blades 654 are rotatable about, for example, a central axis 612. As previously described, fan 652 is driven via a low-pressure shaft 638 and a low-pressure turbine 634. Figure 5 In the embodiment shown, the fan 652 is connected to the low-pressure shaft 638 via a reduction gearbox 655, for example, by means of an indirect drive or gear transmission configuration.

[0066] Furthermore, the fan blades 654 can be arranged at equal intervals around the central axis 612. Each fan blade 654 has a root and a tip, and defines the span between the two.

[0067] Furthermore, each fan blade 654 defines a central blade axis 656. In this embodiment, each fan blade 654 of the fan 652 can rotate about its corresponding central blade axis 656, for example, rotating synchronously. One or more starters 658 are provided to achieve this rotational action, which can be used to change the pitch angle of the fan blade 654 about its corresponding central blade axis 656.

[0068] Fan section 650 further includes a fan guide vane array 660, which includes fan guide vanes 662 arranged around a central axis 612. Figure 5 (Only one is shown). In this embodiment, the fan guide vane 662 cannot rotate about the central axis 612. Each fan guide vane 662 has a root and a tip, and defines an extension between the two. The fan guide vane 662 can be as follows: Figure 5 The structure shown is without a shield, or alternatively, a shield may be provided, for example, by means of an annular shield disposed outward along the radial direction R from the tip of the fan guide vane 662, or attached to the fan guide vane 662.

[0069] Each fan guide vane 662 defines a central blade axis 664. In this embodiment, each fan guide vane 662 of the fan guide vane array 660 is rotatable about its corresponding central blade axis 664, for example, rotating synchronously. One or more starters 666 are provided to achieve this rotational action, thereby allowing for changes in the pitch angle of the fan guide vane 662 about its corresponding central blade axis 664. In other embodiments, each fan guide vane 662 may be fixed and its pitch cannot be changed about its central blade axis 664. The fan guide vanes 662 are mounted on a fan shroud 670.

[0070] like Figure 5 As shown, in addition to the ductless fan 652, a ducted fan 684 is provided behind the fan 652. Thus, the engine 600 simultaneously includes both a ducted fan and a ductless fan, both generating thrust through air movement, and the airflow does not need to pass through at least a portion of the turbine 620 (e.g., in the illustrated embodiment, the airflow does not need to pass through the high-pressure compressor 628 and the combustion chamber). The ducted fan can rotate about approximately the same axis as the fan blades 654. In the illustrated embodiment, the ducted fan 684 is driven by a low-pressure turbine 634 (e.g., connected to a low-pressure shaft 638). As previously mentioned, in the illustrated embodiment, the fan 652 can be referred to as the main fan, and the ducted fan 684 as the secondary fan. It should be understood that "main" and "secondary" are merely convenient terms and do not represent importance, power, or other similar meanings.

[0071] The duct fan 684 includes multiple fan blades ( Figure 5 (Not separately marked). The fan blades of the duct fan 684 can be arranged at equal intervals around the central axis 612. Each blade of the duct fan 684 has a root and a tip, and defines the span between the two.

[0072] The fan shroud 670 annularly covers at least a portion of the core shroud 622 and is located substantially outside at least a portion of the core shroud 622 in the radial direction R. Specifically, the downstream section of the fan shroud 670 covers the front portion of the core shroud 622 to define a fan flow path or fan duct 672. Accordingly, the fan flow path or fan duct 672 can be understood as constituting at least a portion of the third airflow of the engine 600.

[0073] Incoming airflow enters fan duct 672 through fan duct inlet 676 and exits through fan exhaust nozzle 678, thereby generating propulsive thrust. Fan duct 672 is an annular duct, located generally outside the core duct 642 in the radial direction R. Fan fairing 670 and core fairing 622 are interconnected and supported by a plurality of fixed struts 674 that extend substantially radially and are spaced apart circumferentially. Figure 5(Only one support is shown in the diagram). Each fixed support 674 may have an aerodynamic shape to guide airflow passing through it. In addition to the fixed support 674, other supports may be used to connect to and support the fan fairing 670 and / or the core fairing 622. In many embodiments, the fan duct 672 and the core duct 642 may extend at least partially together (generally axially) on both sides (e.g., radially) of the core fairing 622. For example, both the fan duct 672 and the core duct 642 may extend directly from the leading edge 644 of the core fairing 622 and extend generally axially together on both radially sides of the core fairing.

[0074] Engine 600 further defines or includes inlet duct 680. Inlet duct 680 extends between engine inlet 682 and core inlet 624 / fan duct inlet 676. Engine inlet 682 is generally defined at the front end of fan cowl 670 and is located axially between fan 652 and fan guide vane array 660. Inlet duct 680 is an annular duct located radially inside fan cowl 670. Air flowing downstream along inlet duct 680 is not necessarily equally split into core duct 642 and fan duct 672 by the diffuser or leading edge 644 of core cowl 622. The width of inlet duct 680 in the radial direction R is greater than that of core duct 642 and also greater than that of fan duct 672 in the radial direction R.

[0075] During operation of engine 600 under one operating condition, engine 600 generates a total thrust FnTotal. This operating condition may be engine 600 operating at rated speed under standard atmospheric conditions. The total thrust is the sum of a first airflow thrust Fn1s (e.g., the main fan thrust generated by airflow driven by fan 652 flowing through the outside of fan shroud 670 and core shroud 622), a third airflow thrust Fn3s (e.g., the thrust generated by airflow flowing through fan duct 672 and discharged by fan exhaust nozzle 678, which is at least partially generated by duct fan 684), and a second airflow thrust Fn2s (e.g., the thrust generated by airflow flowing through core duct 642 and discharged by turbine exhaust nozzle 640).

[0076] It is worth noting that, in the illustrated embodiment, the engine 600 is provided with one or more structures for improving the efficiency of the third airflow thrust Fn3s. Specifically, the engine 600 further includes an inlet guide vane array 686, which is arranged within the inlet duct 680, upstream of the duct fan 684 and downstream of the engine inlet 682. The inlet guide vane array 686 is arranged around a central axis 612. In this embodiment, the inlet guide vanes 686 cannot rotate about the central axis 612. Each inlet guide vane 686 defines a central blade axis (not labeled for clarity) and can rotate about its corresponding central blade axis, for example, rotating synchronously with each other. One or more starters 668 are provided to achieve this rotation, thereby allowing for changes in the pitch angle of the inlet guide vane 686 about its corresponding central blade axis. In other embodiments, each inlet guide vane 686 may be fixed and its pitch cannot be changed about its central blade axis.

[0077] Furthermore, the engine 600 further provides an outlet guide vane array 690 downstream of the duct fan 684 and upstream of the fan duct inlet 676. Similar to the inlet guide vane array 686, the outlet guide vane array 690 cannot rotate about the central axis 612. In the illustrated embodiment, unlike the inlet guide vane array 686, the outlet guide vane array 690 is configured as a fixed-pitch outlet guide vane.

[0078] Furthermore, it should be understood that in the illustrated embodiment, the fan exhaust nozzle 678 of the fan duct 672 is further configured as a variable geometry exhaust nozzle. In this way, the engine 600 includes one or more starters 692 for adjusting the variable geometry exhaust nozzle. For example, the variable geometry exhaust nozzle may be configured to change the total cross-sectional area (e.g., the area of ​​the nozzle in a plane perpendicular to the central axis 612) to adjust the magnitude of the thrust generated according to one or more engine operating conditions (e.g., the temperature, pressure, mass flow rate, etc. of the airflow flowing through the fan duct 672). A fixed geometry exhaust nozzle may also be used.

[0079] The combination of the inlet guide vane array 686 located upstream of the duct fan 684, the outlet guide vane array 690 located downstream of the duct fan 684, and the fan exhaust nozzle 678 enables the engine to generate third airflow thrust Fn3s more efficiently under one or more operating conditions. Furthermore, by implementing a variable geometry design for the inlet guide vanes 686 and the fan exhaust nozzle 678, the engine 600 can efficiently generate third airflow thrust Fn3s over a wide range of operating conditions, including the takeoff and climb phases (which typically require the engine's total thrust FnTotal to reach its maximum) and the cruise phase (which typically requires a lower engine total thrust FnTotal).

[0080] Continue to refer to Figure 5The air flowing through the fan duct 672 is at a relatively low temperature, for example, lower than the temperature of one or more fluids, such as the air used within the turbine 620. In this way, one or more heat exchangers 699 can be provided to form a heat exchange connection with the fan duct 672. For example, one or more heat exchangers 699 can be arranged inside the fan duct 672, using the air flowing through the fan duct 672 as a cooling medium to cool one or more fluids from the core engine, such as compressor exhaust, oil, or fuel.

[0081] Figure 5 The diagram illustrates various sensors. Measurements from these sensors are used during ground and / or flight testing, as described elsewhere in this document. All of these sensors are connected to the controller.

[0082] For example, an engine control torque sensor 602 is coupled to the low-pressure shaft 638 to measure torque. A first pressure sensor 604 (measuring total pressure), a second pressure sensor 606 (measuring static pressure), and a first temperature sensor 608 (measuring total temperature) are arranged at the outlet of the fan exhaust nozzle 678. A third pressure sensor 610 (measuring total pressure) and a second temperature sensor 613 (measuring total temperature) are arranged at the turbine exhaust nozzle 640. A fourth pressure sensor 615 (measuring total pressure), a fifth pressure sensor 617 (measuring static pressure), and a third temperature sensor 618 (measuring total temperature) are arranged at the engine inlet 682. It should be understood that other sensors may also be arranged in other locations, and the sensors may be of the types described above or other types.

[0083] Shortening climb time and allowing aircraft to spend more time in the fuel-efficient cruise phase can improve fuel efficiency. However, excessive thrust during the climb phase can lead to engine overload, negating the fuel savings from shortened climb times. For example, conventional aircraft cruising at Mach 0.5 and above typically keep their thrust ratio below 1.1 to avoid engine overload and excessive fuel consumption during climb. The method and system described in this paper provide an improved engine thrust control scheme specifically for HARW aircraft. By adjusting the engine thrust ratio for such aircraft, the climb phase can be shortened without engine overload during the ToC period, thereby significantly improving the overall fuel efficiency of HARW aircraft.

[0084] Further aspects of this disclosure are given by the subject of the following clauses:

[0085] A method for operating an aircraft engine. The method includes: controlling the operation of one or more engines of an aircraft during a period including a takeoff phase, a climb phase, and a cruise phase, wherein the climb phase includes a ToC (ToC) period prior to transitioning to the cruise phase; controlling the one or more engines to output takeoff thrust during the takeoff phase, the takeoff thrust corresponding to the maximum thrust during the takeoff phase; controlling the one or more engines to output ToC thrust during the ToC period, the ToC thrust corresponding to the average thrust during the ToC period; and controlling the one or more engines such that the aircraft cruises at Mach 0.5 or higher; wherein the one or more engines are controlled such that a thrust ratio is between 1.1 and 1.6, the thrust ratio being defined as the ratio of the ToC thrust normalized to ToC environmental static pressure to the takeoff thrust normalized to takeoff environmental static pressure.

[0086] According to any of the preceding methods, the engine is an open rotor engine, which includes a rotor and stator blades located behind the rotor.

[0087] The method according to any of the preceding items, wherein the engine is an open rotor engine, and the open rotor engine includes counter-rotating rotors.

[0088] The method according to any of the preceding items, wherein the engine includes an annular core inlet.

[0089] The method described in any of the preceding items includes an aircraft comprising two or more engines.

[0090] The method described in any of the preceding items, wherein it ends at an altitude of 28,000 feet.

[0091] According to the method described in any of the preceding items, the ToC period includes the last two minutes of the climb phase.

[0092] According to the method described in any of the preceding items, the ToC period includes the last thousand or two thousand feet of altitude before reaching cruise altitude and entering the cruise phase.

[0093] According to any of the preceding descriptions, one or more engines are coupled to the wings of an aircraft, the aspect ratio of which is designed to be greater than 10.

[0094] The method described according to any of the foregoing items further includes operating the aircraft to reach the ToC in less than 40 minutes.

[0095] According to the method described in any of the preceding items, each wing of the aircraft includes a folding wingtip.

[0096] According to the method described in any of the preceding items, the truss connects each wing of the aircraft to the fuselage.

[0097] The method described in any of the preceding items, wherein the engine is attached to the fuselage, and the fuselage is designed to carry 270 or fewer passengers.

[0098] The method according to any of the preceding items further includes controlling the thrust produced by each of one or more engines to be less than 50,000 pounds.

[0099] According to the method described in any of the preceding items, each of one or more engines is configured to have a bypass ratio of 15 to 100.

[0100] According to the method described in any of the preceding items, each of one or more engines is configured to have a fan pressure ratio (FPR) of 1.05 to 1.5.

[0101] A method of operating an aircraft engine. The method includes: controlling the operation of one or more engines of an aircraft during a flight mission including a takeoff phase, a climb phase, and a cruise phase, the climb phase including a ToC (ToC) period prior to transitioning into the cruise phase; controlling the one or more engines to output takeoff thrust during the takeoff phase, the takeoff thrust corresponding to the maximum thrust during the takeoff phase; and controlling the one or more engines to output ToC thrust during the ToC period, the ToC thrust corresponding to the average thrust during the ToC period, and the ToC period ending at an altitude above 28,000 feet; wherein the one or more engines are operated such that a thrust ratio is between 1.1 and 1.6, the thrust ratio being defined as the ratio of the ToC thrust normalized to ToC environmental static pressure to the takeoff thrust normalized to takeoff environmental static pressure.

[0102] According to any of the preceding methods, the engine is an open rotor engine, which includes a rotor and stator blades located behind the rotor.

[0103] The method according to any of the preceding items, wherein the engine is an open rotor engine, and the open rotor engine includes counter-rotating rotors.

[0104] The method according to any of the preceding items, wherein the engine includes an annular core inlet.

[0105] The method described in any of the preceding items includes an aircraft comprising two or more engines.

[0106] The method described according to any of the preceding items further includes controlling one or more engines to enable the aircraft to cruise at Mach 0.5 or higher.

[0107] According to the method described in any of the preceding items, the ToC period includes the last two minutes of the climb phase.

[0108] According to the method described in any of the preceding items, the ToC period includes the last thousand or two thousand feet of altitude before reaching cruise altitude and entering the cruise phase.

[0109] According to any of the preceding descriptions, one or more engines are coupled to the wings of an aircraft, the aspect ratio of which is designed to be greater than 10.

[0110] The method described according to any of the foregoing items further includes operating the aircraft to reach the ToC in less than 40 minutes.

[0111] According to the method described in any of the preceding items, each wing of the aircraft includes a folding wingtip.

[0112] According to the method described in any of the preceding items, the truss connects each wing of the aircraft to the fuselage.

[0113] The method described in any of the preceding items, wherein the engine is attached to the fuselage, and the fuselage is designed to carry 270 or fewer passengers.

[0114] The method according to any of the preceding items further includes controlling the thrust produced by each of one or more engines to be less than 50,000 pounds.

[0115] According to the method described in any of the preceding items, each of one or more engines is configured to have a bypass ratio of 15 to 100.

[0116] According to the method described in any of the preceding items, each of one or more engines is configured to have a fan pressure ratio (FPR) of 1.05 to 1.5.

[0117] An aircraft engine. The engine includes: one or more thrust actuators; and an engine controller coupled to the one or more thrust actuators; wherein the engine controller is configured to control the one or more thrust actuators during takeoff, climb, and cruise phases based on instructions received from a flight controller, wherein the climb phase ends at a peak climb (ToC) period at an altitude above 28,000 feet; and wherein the engine controller controls the thrust generated by the engine such that the ratio of the engine peak climb (ToC) thrust normalized to the ambient static pressure to the takeoff (TO) thrust normalized to the ambient static pressure is in the range of 1.1 to 1.6.

[0118] The engine described in any of the preceding items, wherein the thrust produced by the engine is less than 50,000 pounds.

[0119] The engine described in any of the preceding items is configured to have a bypass ratio of 15 to 100.

[0120] The engine according to any of the preceding items, wherein the engine is configured to have a fan pressure ratio (FPR) of 1.05 to 1.5.

[0121] An aircraft. The aircraft includes: an aircraft engine as described in any of the preceding items; a fuselage; and a pair of wings.

[0122] The aircraft described in any of the preceding items, wherein the wing dimensions of the aircraft are designed with an aspect ratio greater than 10.

[0123] According to any of the preceding items, the engine has a travel time to ToC of less than 40 minutes.

[0124] The engine described in any of the preceding items, wherein each of the pair of wings of the aircraft includes folding wingtips.

[0125] The engine is as described in any of the preceding items, wherein the truss connects each wing of the aircraft to the fuselage.

[0126] A computer-readable storage medium storing instructions, which, when executed by a processor, cause the processor to perform a method for operating engines in an aircraft, the method comprising: controlling the operation of one or more engines of the aircraft during a period including a takeoff phase, a climb phase, and a cruise phase, the climb phase including a peak climb (ToC) period before transitioning to the cruise phase; controlling the one or more engines to output takeoff thrust during the takeoff phase, the takeoff thrust corresponding to the maximum thrust during the takeoff phase; controlling the one or more engines to output ToC thrust during the ToC period, the ToC thrust corresponding to the average thrust during the ToC period; and controlling the one or more engines such that the aircraft cruises at Mach 0.5 or higher; wherein the one or more engines are controlled such that a thrust ratio is between 1.1 and 1.6, the thrust ratio being defined as the ratio of the ToC thrust normalized to the ToC environmental static pressure to the takeoff thrust normalized to the takeoff environmental static pressure.

[0127] A computer-readable storage medium storing instructions, when executed by a processor, causing the processor to perform a method for operating engines in an aircraft, the method comprising: controlling the operation of one or more engines of the aircraft during a flight mission including a takeoff phase, a climb phase, and a cruise phase, the climb phase including a peak climb (ToC) period prior to transitioning into the cruise phase; controlling the one or more engines to output takeoff thrust during the takeoff phase, the takeoff thrust corresponding to the maximum thrust during the takeoff phase; and controlling the one or more engines to output ToC thrust during the ToC period, the ToC thrust corresponding to the average thrust during the ToC period, and the ToC period ending at an altitude above 28,000 feet; wherein the one or more engines are operated such that a thrust ratio is between 1.1 and 1.6, the thrust ratio being defined as the ratio of the ToC thrust normalized to ToC environmental static pressure to the takeoff thrust normalized to takeoff environmental static pressure.

[0128] This written description discloses the disclosure by way of example, including the best implementation, and is also intended to enable those skilled in the art to implement the disclosure, including making and using any apparatus or system, and implementing any incorporated method. The scope of protection of this disclosure is defined by the claims and may cover other examples readily conceived by those skilled in the art. Such other examples are intended to fall within the scope of the claims if they contain equivalent structural elements that are not materially different from those stated in the claims, or contain structural elements that are not different from those stated in the claims.

Claims

1. A method for operating an aircraft engine, characterized in that, The method includes: During a mission that includes a takeoff phase, a climb phase, and a cruise phase, the operation of one or more engines of the aircraft is controlled, the climb phase including the peak climb (ToC) period before transitioning to the cruise phase; During the takeoff phase, the one or more engines are controlled to output takeoff thrust, the takeoff thrust corresponding to the maximum thrust during the takeoff phase; During the ToC period, control the one or more engines to output ToC thrust, the ToC thrust corresponding to the average thrust during the ToC period; and Control the one or more engines to enable the aircraft to cruise at Mach 0.5 or higher; Specifically, the one or more engines are controlled such that the thrust ratio is between 1.1 and 1.6, and the thrust ratio is defined as the ratio of the ToC thrust normalized to the ToC environmental static pressure to the takeoff thrust normalized to the takeoff environmental static pressure.

2. The method according to claim 1, characterized in that, in, The one or more engines include an open rotor engine, which includes a rotor and stator blades located behind the rotor.

3. The method according to claim 1, characterized in that, in, The one or more engines include an open rotary engine, which includes counter-rotating rotors.

4. The method according to claim 1, characterized in that, in, The one or more engines include engines with annular core inlets.

5. The method according to claim 1, characterized in that, in, The aircraft includes two or more engines.

6. The method according to claim 1, characterized in that, in, The ToC period ends at an altitude of over 28,000 ft.

7. The method according to claim 1, characterized in that, in, The ToC period includes the last two minutes of the climb phase.

8. The method according to claim 1, characterized in that, in, The ToC period occurs at an altitude within two thousand feet below the cruising altitude of the cruise phase.

9. The method according to claim 1, characterized in that, in, The one or more engines are connected to the wings of the aircraft, the wings being sized to have an aspect ratio greater than 10, the aspect ratio being the ratio of the square of the wingspan to the projected area of ​​the wing.

10. The method according to claim 1, characterized in that, in, The one or more engines include pipeline engines.

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