• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    gas turbine engine The present invention relates to a gas turbine engine for an aircraft, the gas turbine engine (10) comprising: a central core (11) comprising a turbine (19), a compressor (14) and a central shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the central core, the fan comprising a plurality of fan blades; and a nacelle (21) surrounding the central core (11) and defining a bypass duct (22) and a bypass exhaust nozzle (18), wherein the gas turbine engine (10) is configured so that a first velocity ratio between an exhaust flow axial velocity from the turbine and a fully expanding exhaust flow axial velocity from the bypass exhaust nozzle is greater than about 0.655 in take-off maximum thrust conditions. [Figure 1]
    公开号:FR3096407A1
    申请号:FR2004461
    申请日:2020-05-05
    公开日:2020-11-27
    发明作者:Craig W Bemment
    申请人:Rolls Royce PLC；
    IPC主号:
    专利说明:

    [0001] Description Title of the invention: Gas turbine engine
    [0001] The present description relates to a gas turbine engine for an aircraft.
    [0002] [0002] By-flow gas turbine engines for propelling aircraft have many design factors which affect the overall efficiency and the power or thrust output.
    [0003] [0003] A high propulsion efficiency for a geared gas turbine engine is obtained by virtue of a high mass flow rate through the engine.
    [0004] [0004] A general objective of geared gas turbine engines in particular, as the fan increases in diameter, is to be able to design the engine with high propulsion efficiency, and therefore low specific combustion of fuel, as well as efficiently integrating the engine into an aircraft with a minimum installation penalty, that is to say with the minimum modifications necessary for the overall design of the aircraft.
    [0005] According to a first aspect, a gas turbine engine for an aircraft is provided, comprising:
    [0006] A central core comprising a turbine, a compressor, and a central shaft connecting the turbine to the compressor;
    [0007] [0007] a fan located upstream of the central core, the fan comprising a plurality of fan blades; and
    [0008] A nacelle surrounding the central core and defining a bypass duct and a bypass exhaust nozzle,
    [0009] [0009] wherein the gas turbine engine is configured such that a first speed ratio between an axial speed of exhaust flow from the turbine and an axial speed of fully expanding exhaust flow at from the bypass exhaust nozzle is greater than about 0.655 under conditions of maximum take-off thrust.
    [0010] Maximum take-off thrust (MTO) conditions can be defined as engine operation under pressure and temperature conditions at sea level of the International Standard Atmosphere (ISA) +15 ° C at thrust maximum take-off at the end of the runway, which is typically defined at an aircraft speed of about 0.25 Mn, or between about 0.24 and 0.27 Mn.
    [0011] A gas turbine engine according to the invention allows a reduced turbine exhaust area through the use of a higher output speed, which allows the turbine to be operated at a higher rotational speed and cc made with reduced overall diameter.
    [0012] [0012] The first speed ratio can, in certain examples, be greater than approximately 0.69 under conditions of maximum take-off thrust.
    [0013] [0013] The engine can be further configured such that a second speed ratio between the axial speed of fully expanding exhaust flow from the bypass exhaust nozzle under conditions of maximum take-off thrust and under cruising conditions is less than about 0.82, and possibly greater than about 0.7.
    [0014] A first engine bypass ratio under maximum take-off thrust conditions divided by a second engine bypass ratio under cruising conditions may be less than approximately 0.82.
    [0015] A first specific engine thrust under maximum conditions divided by a second specific thrust under cruising conditions may be less than about 0.82.
    [0016] The first gear ratio can be less than about 1.1 or 1.0.
    [0017] An engine bypass ratio may be greater than about 10 at cruising conditions, for example between about 10 and about 20 at cruising conditions.
    [0018] The gas turbine engine may include a gearbox which receives an input from the central shaft and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the central shaft.
    [0019] According to a second aspect, a method of operating a gas turbine engine on an aircraft is provided, the gas turbine engine comprising:
    [0020] A central core comprising a turbine, a compressor, and a central shaft connecting the turbine to the compressor;
    [0021] A fan located upstream of the central core, the fan comprising a plurality of fan blades; and
    [0022] A nacelle surrounding the central core and defining a bypass duct and a bypass exhaust nozzle,
    [0023] [0023] wherein the method comprises operating the gas turbine engine at maximum take-off thrust conditions such that a first speed ratio between an axial speed of exhaust flow from the turbine and a speed Axial of fully expanding exhaust flow from the bypass exhaust nozzle is greater than about 0.655.
    [0024] The optional and advantageous characteristics described above in relation to the first aspect can also be applied to the method according to the second aspect.
    [0025] The cruising conditions can be defined as a forward Mach number of between 0.7 and 0.9 at an altitude of between 10,000 m and 15,000 m.
    [0026] As indicated elsewhere in this document, this description may relate to a gas turbine engine.
    [0027] Arrangements of the present description may be particularly, although not exclusively, advantageous for blowers which are driven through a gearbox.
    [0028] The gearbox may be an epicyclic gearbox comprising an input sun wheel connected to the central shaft, a plurality of satellites connected by a carrier arm and an outer ring, the fan being connected to the arm. carrier.
    [0029] The gas turbine engine as described and / or claimed here can have any suitable general architecture.
    [0030] In such an arrangement, the second compressor can be positioned axially downstream of the first compressor.
    [0031] [0031] The gearbox can be arranged to be driven pal- the central shaft which is configured to rotate (for example in use) at the lowest rotational speed (for example the first central shaft in the example. above).
    [0032] The gearbox is a reduction box (in that the outlet to the fan has a lower rotational speed than the inlet from the central shaft).
    [0033] In any gas turbine engine as described and / or claimed here, a combustion chamber may be provided axially downstream of the fan and of the compressor (s).
    [0034] The compressor (s) (for example the first compressor and the second compressor as described above) can comprise any number of stages, for example several stages.
    [0035] The turbine (s) (for example the first turbine and the second turbine as described above) can comprise any number of stages, for example several stages.
    [0036] [0036] Each fan blade can be defined as having a radial extent 6 extending from a foot (or hub) at a radially internal gas-washed location, or position of 0% extent, up to a tip at a 100% stretch position.
    [0037] The radius of the fan can be measured between the center line of the engine and the tip of a fan blade at its leading edge.
    [0038] The speed of rotation of the fan can vary in use.
    [0039] When using the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation.
    [0040] Gas turbine engines according to the present description can have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core in cruising conditions.
    [0041] The overall pressure ratio of a gas turbine engine as described and / or claimed here can be defined as the ratio of the total pressure upstream of the fan to the total pressure at the outlet. of the high pressure compressor (before entering the combustion chamber).
    [0042] [0042] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow rate through the engine.
    [0043] [0043] A gas turbine engine as described and / or claimed herein can have any desired maximum thrust.
    [0044] In use, the temperature of the flow at the inlet of the high pressure turbine can be particularly high.
    [0045] [0045] A fan blade and / or profile portion described and / or claimed herein can be made from any suitable material or any suitable combination of materials.
    [0046] A fan as described and / or claimed herein may include a central portion, from which the fan blades may extend, for example in a radial direction.
    [0047] The gas turbine engines described and / or claimed here may or may not be provided with a variable area nozzle (VAN).
    [0048] The fan of a gas turbine as described and / or claimed herein can have any number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades. .
    [0049] As used here, the term cruising conditions has a conventional meaning and will be easily understood by those skilled in the art.
    [0050] In other words, for a given gas turbine engine for an aircraft, the cruising conditions are defined as the engine operating point which provides a specified thrust (required to provide, in combination with any other engine of the aircraft, continuous operation of the aircraft on which it is intended to be installed at a given Mach number at mid-cruise) under atmospheric conditions at mid-cruise (defined by the international standard atmosphere according to ISO 2533 at mid-cruise altitude).
    [0051] By way of example only, the forward speed in cruising conditions can be any point in the range of Mach 0.7 to 0.9, for example 0.75 to 0, 85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example from l of the order of Mach 0.8, of the order of Mach 0.85 or in the range of 0.8 to 0.85.
    [0052] By way of example only, the cruising conditions may correspond to standard atmospheric conditions (according to the international standard atmosphere, ISA) at an altitude in the range of 10,000 m to 15,000 ni, for example in the range 10,000 m to 12,000 m, for example in the range 10,400 ni to 11,600 m (about 38,000 ft), for example in the range 10,500 ni to 11,500 m, for example in the range 10,600 m to 11,400 m, for example in the range 10,700 m (approx. 35,000 ft) to 11,300 m, for example in the range 10,800 m to 11,200 m, for example in the range 10 900 m to 11,100 m, for example of the order of 11,000 m.
    [0053] By way of example only, the cruising conditions may correspond to an engine operating point which provides a known required level of thrust (for example a value in the range of 30 kN to 35 kN) to a number of Mach 0.8 and standard atmospheric conditions (according to the international standard atmosphere) at an altitude of 38,000 ft (11,582 m).
    [0054] In use, a gas turbine engine described and / or claimed here can operate under the cruising conditions defined here.
    [0055] In one aspect, there is provided an aircraft comprising a gas turbine engine as described and / or claimed herein.
    [0056] [0056] According to one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein.
    [0057] Those skilled in the art will understand that, except mutual exclusivity, a characteristic or a parameter described in relation to any one of the above aspects can be applied to any other aspect.
    [0058] Embodiments will now be described by way of example only, with references to the Figures, in which:
    [0059] [0059] [fig.1] is a side sectional view of a gas turbine engine;
    [0060] [0060] [Fig.2] is a close-up side sectional view of a part upstream of a gas turbine engine;
    [0061] [0061] [Fig.3] is a partially cut away view of a gearbox for a gas turbine engine;
    [0062] [0062] [Fig.4] is a schematic drawing of an aircraft having a gas turbine engine mounted thereon;
    [0063] [0063] [fig.5] is a schematic drawing illustrating the concept of a fully expanding jet velocity; and
    [0064] [0064] [Fig.6] is an example of a plot showing the relationship between the Mach number at the bypass exhaust nozzle and the bypass nozzle pressure ratio.
    [0065] Figure 1 illustrates a gas turbine engine 10 having an axis of rotation 9.
    [0066] In use, the central air flow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where additional compression occurs.
    [0067] An exemplary arrangement for a gas turbine engine 10 is illustrated in Figure 2.
    [0068] It should be noted that the terms "low pressure turbine" and "low pressure compressor" can be used to denote the minimum pressure turbine stages and the minimum pressure compressor stages (i.e. n 'not including the blower 23) respectively and / or the turbine and compressor stages interconnected by the interconnection shaft 26 with the lowest rotational speed in the engine (i.e. n' not including the gearbox output shaft which drives the blower 23).
    [0069] The epicyclic gearbox 30 is shown by way of example in greater detail in Figure 3.
    [0070] The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via links 36, the crown 38 being fixed.
    [0071] It is understood that the arrangement illustrated in Figures 2 and 3 is provided by way of example only, and various alternatives are within the scope of the present description.
    [0072] [0072] Accordingly, the present invention extends to a gas turbine engine having any arrangement of gearbox styles (eg star or planetary), support structures, shaft arrangement, etc. entry and exit, and landing locations.
    [0073] [0073] Optionally, the gearbox can drive additional and / or alternative components (for example, the intermediate pressure compressor and / or a precompressor).
    [0074] Other gas turbine engines to which the present description can apply may have alternative configurations.
    [0075] The geometry of the gas turbine engine 10, and its components, is defined by a conventional axial system, comprising an axial direction (which is aligned with the axis of rotation 9), a radial direction (from the bottom to the top in Figure 1), and a circumferential direction (perpendicular to the page in the view of Figure 1).
    [0076] Also referring to Figure 1, Tc is the axial speed of exhaust flow (or cold nozzle speed) from the turbine and VB is the axial speed of exhaust flow from the nozzle d 'bypass exhaust 18.
    [0077] A ratio between the speed of the cold nozzle in total expansion Vc. between maximum take-off thrust conditions and cruise conditions may be less than 0.82.
    [0078] Figure 4 illustrates an example of an aircraft 40 having a gas turbine engine 10 attached to each wing 41a, 41b thereof.
    [0079] Figure 5 illustrates an example of an exhaust nozzle 50 of a gas turbine engine.
    [0080] [0080] Figure 6 is an example of a plot showing the relationship between the Mach number at the bypass exhaust nozzle 18 (see Figure 1) and the bypass nozzle pressure ratio, i.e. the ratio between the total pressure at the bypass exhaust nozzle and ambient pressure.
    [0081] Parameters which can be adjusted to achieve a speed ratio within the required range may include fan blade outlet (particularly at the fan foot) and ESS inlet area.
    [0082] The following table illustrates examples of parameters for two examples of engine, Example 1 being for an engine of relatively low power, or less, and Example 2 for an engine of relatively high power, or greater.
    [0083] [0083] 17 [Tables 1] Parameter Example 1 (small motor) Example 2 (large motor) Blower diameter (cm) 215 320 Total LPT outlet pressure at maximum flow (kPa) 130 130 Maximum LPT outlet mass flow rate (k gis) 50 100 Branch area of LPT rotor (m2) 0.38 or less, e.g. 0.25 to 0.38 0.75 or less, e.g. 0.5 to 0.75 Total inlet pressure ESS to flow maximum (kPa) 140 140 ESS inlet mass flow rate (kg / s) 50 100 ESS inlet rotor area (m2) 0.275 or more, for example 0.27 to 0.3 0.55 or more, for example 0.55 to 0.6 100841 The above parameters relating to the total LPT outlet pressure at maximum flow, the maximum LPT outlet mass flow rate and the final LPT rotor area together determine the output flow speed of the LPT, i.e. the axial velocity of exhaust flow from the turbine.
    [0085] To reduce the input Mach number, the average ESS input radius can be increased, which can be done while maintaining a given ESS input range.
    [0086] It will be understood that the invention is not limited to the embodiments described above and that various modifications and improvements can be made without departing from the concepts described here.
    Unless mutually exclusive, any feature may be employed separately or in combination with other features and the description extends to and includes all combinations and sub-combinations of one or more features described herein.
    19 [Claim 1] [Claim 2] [Claim 3] [Claim 4]
    权利要求:
    Claims (1)
    [0001]
    A gas turbine engine (10) for an aircraft comprising: a central core (11) comprising a turbine (19), a compressor (14) and a central shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the central core, the fan comprising a plurality of fan blades; and a nacelle (21) surrounding the central core (11) and defining a bypass duct (22) and a bypass exhaust nozzle (18), in which the gas turbine engine (10) is configured so that a first speed ratio between an axial speed of exhaust flow from the turbine (19) and an axial speed of fully expanding exhaust flow from the bypass exhaust nozzle (18) is greater than about 0.655 under maximum take-off thrust conditions. A gas turbine engine (10) according to claim 1 comprising a gearbox (30) which receives an input from the central shaft (26) and outputs a drive to the fan so as to drive the fan at a rate. lower rotational speed than the central shaft, in which optionally the gear ratio is in the range of 3.1 to 4.2, optionally 3.2 to 3.8. A gas turbine engine according to claim 2, wherein: the turbine is a first turbine (19), the compressor is a first compressor (14), and the central shaft is a first central shaft (26); the central core further comprises a second turbine (17), a second compressor (15), and a second central shaft (27) connecting the second turbine to the second compressor; and the second turbine, the second compressor and the second central shaft may be arranged to rotate at a higher rotational speed than that of the first central shaft. A method of operating a gas turbine engine (10) on an aircraft (40), the gas turbine engine (10) comprising: a central core (11) including a turbine (19), a compressor (14) and a central shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the central core, the fan comprising a plurality of fan blades; and a pod (21) surrounding the central core (11) and defining a bypass duct (22) and a bypass exhaust nozzle (18), 20 [Claim 5] [Claim 6] [Claim 7] [Claim 8 ] [Claim 9] wherein the method comprises operating the gas turbine engine (10) under conditions of maximum takeoff thrust such that a first speed ratio between an axial speed of exhaust flow from turbine (19) and an axial velocity of fully expanding exhaust flow from the bypass exhaust nozzle (18) is greater than about 0.655. A method according to any one of claim 4, wherein the gas turbine engine (10) comprises a gearbox (30) which receives an input from the center shaft (26) and outputs a drive to. the blower so as to drive the blower at a lower rotational speed than the central shaft, in which optionally the gear ratio is in the range of 3.1 to 4.2, optionally 3.2 to 3.8. A method according to claim 5, wherein: the turbine is a first turbine (19), the compressor is a first compressor (14), and the central shaft is a first central shaft (26); the central core further comprises a second turbine (17), a second compressor (15), and a second central shaft (27) connecting the second turbine to the second compressor; and the second turbine, the second compressor and the second central shaft rotate at a higher rotational speed than that of the first central shaft. A gas turbine engine (10) according to any one of claims 1 to 3, or a method according to any one of claims 4 to 6, wherein, under conditions of maximum takeoff thrust, the first gear ratio is greater than about 0.69 and / or less than about 1.1, possibly less than about 1.0. A gas turbine engine (10) according to any one of claims 1 to 3 or claim 7 when dependent thereon, or a method according to any one of claims 4 to 6 or claim 7 when dependent thereon. depends on these, wherein, a second speed ratio between the axial velocity of fully expanding exhaust flow from the bypass exhaust nozzle under conditions of maximum take-off thrust and under cruise conditions is less than about 0.82, and possibly greater than about 0.7. A gas turbine engine (10) according to any one of claims 1 to 3 or any claim dependent thereon, or a method according to any one of claims 4 to 6 or one 21 [Claim 10 ] Any claim dependent thereon, wherein an engine bypass ratio is in the range of 10 to 20 at cruising conditions, optionally in the range of 13 to 18. A gas turbine engine (10) according to any one of claims 1 to 3 or any claim dependent thereon, or a method according to any one of claims 4 to 6 or any claim dependent thereon, wherein conditions of Maximum takeoff thrust are defined as the engine running at maximum takeoff thrust at ISA sea level pressure and temperature +15 ° C with a fan inlet speed of 0.25 Mn.
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    同族专利:
    公开号 | 公开日
    DE102020113051A1|2020-11-26|
    US20200370512A1|2020-11-26|
    GB201907258D0|2019-07-10|
    CN213510751U|2021-06-22|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    GB201908978D0|2019-06-24|2019-08-07|Rolls Royce Plc|Gas turbine engine transfer efficiency|
    GB201908972D0|2019-06-24|2019-08-07|Rolls Royce Plc|Compression in a gas turbine engine|
    法律状态:
    2021-05-26| PLFP| Fee payment|Year of fee payment: 2 |
    优先权:
    申请号 | 申请日 | 专利标题
    GB1907258.6|2019-05-23|
    GBGB1907258.6A|GB201907258D0|2019-05-23|2019-05-23|Gas turbine engine|
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