METHOD AND SYSTEM FOR REGULATING THE THRUST OF AN AIRCRAFT TURBOMACHINE

Abstract

A method and system control the thrust of an aircraft turbomachine having a high bypass ratio by direct action on a variable-pitch system. The variable-pitch system varies the pitch of the vanes of a stator of a low-pressure compressor for the open-loop control of the thrust of the turbomachine. The method also provides closed-loop control of the pitch of the blades of a propeller based on a rotational speed of the propeller.

Claims

1. A method for regulating the thrust of an aircraft turbomachine (100), for an aircraft turbomachine (100) comprising: a propeller (110) comprising a plurality of blades; a blade variable pitch setting system (112) for varying the pitch of said blades; successively along an engine axle: a low-pressure compressor (120) comprising at least one straightener (121) equipped with a vane variable pitch setting system (123) for changing an orientation of vanes of the straightener (121), a high-pressure compressor (130), a combustion chamber (160), one or more turbines (140, 150); the method being characterised in that it comprises the following steps: (i) varying a pitch setting of vanes of the straightener (121) by means of the vane variable pitch setting system (123) to regulate in an open loop (B0) the thrust of the turbomachine (100); (ii) regulating a pitch setting of blades of the propeller (110) in a closed loop (B2) based on a rotational speed of the propeller (110), by means of the blade variable pitch setting system (112).

2. The method according to the preceding claim, characterised in that it further comprises the following step (0) which precedes the steps (i) and (ii): (0) selecting a steady regime for the aircraft turbomachine (100); the pitch setting of blades of the propeller (110) being regulated at step (ii) based on the steady regime selected at step (0).

3. The method according to the preceding claim, characterised in that the pitch setting of blades of the propeller (110) is regulated at step (ii) so that the rotational speed of the propeller (110) corresponds to a first mathematical function of the pitch setting of vanes of the straightener (121), the first function depending on the steady regime selected at step (0).

4. The method according to the preceding claim, characterised in that the first function is one of: a substantially constant function, a bijective monotonic function, preferably affine.

5. The method according to any of the preceding claims, further comprising the following step: (iii) regulating a fuel flow rate injected into the combustion chamber (160) in a closed loop (B1) based on a ratio of reduced rotational speeds of the high (130) and low (120) pressure compressors.

6. The method according to the preceding claim when dependent on any of claims 2 to 4, characterised in that the flow rate of fuel injected into the combustion chamber (160) is regulated at step (iii) on the basis of the steady regime selected at step (0).

7. The method according to the preceding claim, characterised in that the flow rate of fuel injected into the combustion chamber (160) is regulated at step (iii) so that said ratio corresponds to a second mathematical function of the pitch setting of vanes of the straightener (121), the second function depending on the steady regime selected at step (0).

8. The method according to the preceding claim, characterised in that the second function is one of: a substantially constant function, a bijective monotonic function, preferably affine.

9. The method according to any of the preceding claims, characterised in that the aircraft turbomachine (100) is of a bypass ratio of at least 20.

10. The method according to any one of the preceding claims, characterised in that the aircraft turbomachine (100) comprises a plurality of propellers (110, 110′) equipped with blades, each associated with a blade variable pitch setting system (112, 112′) for varying the pitch of its blades, and in that the step (ii) applies for each of the propellers (110, 110′) by means of the blade variable pitch setting system (112, 112′) associated to it.

11. A system for regulating the thrust of an aircraft turbomachine (100) characterised in that it comprises means for implementing the method according to any of the preceding claims.

12. An aircraft turbomachine (100) comprising: a propeller (110) comprising a plurality of blades; a blade variable pitch setting system (112) for varying the pitch of said blades; successively along an engine axle: a low-pressure compressor (120) comprising at least one straightener (121) equipped with a vane variable pitch setting system (123) for changing an orientation of vanes of the straightener (121), a high-pressure compressor (130), a combustion chamber (160), one or more turbines (140, 150); characterised in that it comprises the regulation system according to the preceding claim.

13. The aircraft turbomachine (100) according to the preceding claim, characterised in that the propeller (110) and the low-pressure compressor (120) are connected by means of a low-pressure shaft (101) and a reducer (111), so that the rotational speeds of the propeller (110) and of the low-pressure compressor (120) are proportional.

14. The aircraft turbomachine (100) according to claim 13, characterised in that it consists of a double-flow turbojet engine, the propeller (110) being ducted.

15. The aircraft turbomachine (100) according to claim 13, characterised in that it consists of an open-rotor thruster, the propeller (110) being un-ducted.

Description

BRIEF DESCRIPTION OF THE FIGURES

[0069] Further characteristics and advantages of the present invention will become apparent from the following detailed description, for the understanding of which reference is made to the attached figures, among which:

[0070] each of FIGS. 1 and 2 illustrates a simplified schematic view of a two-dimensional cross-section of a preferred embodiment of an aircraft turbomachine on which it is intended to integrate the regulation system according to the invention;

[0071] each of FIGS. 3 and 4 illustrate an operating field of a low-pressure compressor of an aircraft turbomachine;

[0072] FIG. 5 illustrates schematically a preferred embodiment of the regulation method according to the invention.

[0073] The drawings in the figures are not to scale. Generally, similar elements are denoted by similar references in the figures. In the scope of this document, the same or similar elements may have the same references. Furthermore, the presence of reference numbers or letters in the drawings cannot be considered as limiting, even when these numbers or letters are indicated in the claims.

DETAILED DESCRIPTION OF PARTICULAR EMBODIMENTS OF THE INVENTION

[0074] In the case of this document, “a propeller of an aircraft turbomachine” refers to both a ducted propeller and an un-ducted propeller. As noted in the prior art, a large, variable pitch setting “fan” is currently considered as a “ducted propeller”. This portion of the text provides a detailed description of preferred embodiments of the present invention. The latter is described with particular embodiments and references to figures but the invention is not limited by them. The drawings and/or figures described below are schematic only and are not limiting.

[0075] References are shown in some of these figures as abstract geometrical reference frames primarily to quantify and/or visualise properties of embodiments of the invention. The reference Z is, for example, usually the “engine axle” of the aircraft turbomachine. This is directed from “upstream” to “downstream”. The stages of the compressors and the turbines of the aircraft turbomachine are stacked essentially along this engine axle. The terms “inlet” and “outlet” of a compressor refer to the upstream and downstream ends of the compressor respectively. In the context of this document, reference is made to the “axial”, “circumferential” and “radial” directions, which correspond preferentially and respectively to directions parallel to the engine axle, essentially circular around the engine axle, and direction perpendicular to the engine axle. Reference frames in FIGS. 1 and 2 illustrate these directions (equipped with an orientation) noted as Z, R and C respectively. The direction and the orientation of the engine axle and the vector Z correspond. The terms “axially” and “radially” are derived from the terms “axial” and “radial” respectively with a similar preferred meaning. The terms “circumferential” and “radial” preferably refer to a polar coordinate system known to a person skilled in the art in each plane perpendicular to the engine axle.

[0076] FIGS. 1 and 2 illustrate cross-sections of two axial aircraft turbomachines 100 on which it is intended to regulate the thrust by the method according to the present invention. Each of these turbomachines 100 comprises a propeller 110 equipped with a plurality of blades, associated with a blade variable pitch setting system 112 for varying the pitch of the blades, and successively along the engine axle (parallel to the axis Z), a low-pressure compressor 120, a high-pressure compressor 130, a combustion chamber 160, a high-pressure turbine 140 and a low-pressure turbine 150. These elements are known to a person skilled in the art. In operation, the mechanical power of the low-pressure 150 and high-pressure 140 turbines is transmitted by means of the low-pressure 101 and high-pressure 102 shafts to the low-pressure 120 and high-pressure 130 compressors respectively, as well as to the propeller 110 via the low-pressure 101 shaft by means of an interposed reducer 111, so that the rotational speeds of the propeller 110 and the low-pressure 120 compressor are proportional. Although not systematically referenced in FIGS. 1 and 2, each compressor 120, 130 and each turbine 140, 150 comprises at least one stage, each such stage comprising a stationary blading and a movable blading adapted to be rotated about the engine axle. For the low-pressure compressor 120, these stationary and movable bladings are denoted 121-121′ and 122 respectively. The stationary bladings of the compressors 120, 130 are generally referred to as “straighteners”. The movable and stationary bladings of the stages of such a compressor 120, 130 are alternated along the engine axle. Parameters such as the dimensions and the surface geometry of the vanes are determined so that the operating conditions of each stage are adapted to those of the upstream and/or downstream stages along the engine axle. In particular, the movable bladings provide an energy by increasing the relative speed of the flowing of an air flow passing through the compressor 120, 130, while the straighteners bring the flowing back parallel to the engine axle while increasing the pressure and decreasing the absolute speed of the flowing. Each stationary vane has a given aerodynamic profile and a pitch setting angle relative to the engine axle to impose a flowing orientation. The low-pressure 120 and high-pressure 130 compressors thus allow to synergistically suck in and compress air so as to bring it to suitable speed, pressure and temperature adapted at the inlet of the combustion chamber 160. These concepts are known to a person skilled in the art. One or more of the straighteners of the compressors 120, 130 are equipped with a vane variable pitch setting system for varying the pitch of their vanes to optimise the flowing of the air flow between stages of these compressors 120, 130. Similarly, the blade variable pitch setting system 112 for varying the pitch of blades coupled to the propeller 110 allows the orientation of the profile of the blades and therefore the pitch of this propeller 110 to be modified, in order to obtain good operation of the aircraft turbomachine 100 according to the different phases of flight encountered. As is well known to a person skilled in the art, each of the above-mentioned vane variable pitch setting systems for varying vanes of straighteners preferably comprises an annulus (referred to as “VSV”) mechanically coupled to the vanes and adjusted circumferentially and externally around an associated casing of the compressor 120, 130, and a displacement unit, typically at least one jack, for displacing this annulus, and accordingly modifying in a synchronised manner the pitch setting of the vanes of the straightener. According to FIGS. 1 and 2, the low-pressure compressor 120 comprises at least one straightener 121 (hereinafter simply referred to as “the straightener 121”) equipped with such a variable pitch setting system referenced by 123. This straightener 121 is preferably upstream of the low-pressure compressor 120. The representation in FIG. 1 is not limiting to the case where one or more of the other straighteners 121′ is also equipped with such a variable pitch setting system. For each of the aircraft turbomachines 100 shown in FIGS. 1 and 2, a hot air flow, referred to as primary flow 106, passes through the aircraft turbomachine 100 axially, thereby feeding the combustion chamber 160, while a cold air flow, referred to as secondary flow 107, is primarily intended to generate a thrust reaction necessary for the flight of the aircraft. This secondary flow 107 is accelerated by the propeller 110 without passing through the combustion chamber 160. The ratio of the secondary 107 and main 106 mass flows is referred to as bypass ratio. In the case of the invention, it is preferably equal to 20 or more because the propeller 110 is preferably large.

[0077] The propeller 110 of the aircraft turbomachine 100 shown in FIG. 1 is ducted and typically referred to as fan. It allows the generation of primary 106 and secondary 107 flows upstream of the low-pressure compressor 120. This aircraft turbomachine 100 is preferably a UHBR type turbojet engine.

[0078] The aircraft turbomachine 100 shown in FIG. 2 is a CROR thruster of the class of the “open rotor” thrusters. The propeller 110 is un-ducted. This thruster also comprises a second, un-ducted propeller 110′ equipped with blades. In the scope of the embodiment, these propellers 110, 110′ are counter-rotating. The case where one of these propellers 110, 110′ is stationary is not excluded from the invention. As these propellers are not ducted, they can be easily sized to accelerate a very large secondary mass flow 107. Each propeller 110, 110′ is coupled to a blade variable pitch setting system 112, 112′ for varying the pitch of its blades and to the low-pressure shaft 101 by substantially radially extending mechanical arms 113, 113′. The method according to the invention is preferably applied to each of these propellers 110, 110′ with its associated variable pitch setting 112, 112′ of blades.

[0079] This invention proposes to regulate the thrust of aircraft turbomachines 100 as illustrated in FIGS. 1 and 2 by means of the regulation method according to the invention. This is illustrated in FIG. 5, and commented on below for a preferred execution in steady regime. The method can be implemented by a regulation system as introduced in the disclosure. In accordance with the steps (i) and (ii), the method proposes: [0080] an open loop B0 for regulating the thrust (directly operable by a thrust lever 1, e.g. by a pilot) by means of the vane variable pitch setting system 123 based on the pitch setting of the vanes of the straightener 121; [0081] a closed loop B2 for regulating the pitch setting of the blades of the propeller 110 by means of the blade variable pitch setting system 112 based on a rotational speed of the propeller 110.

[0082] This closed loop B2 is a function of the steady regime selected on the basis of the appropriate device 2 in the aircraft turbomachine, in the sense that this regime is preferably associated in advance with a rotational speed of the propeller 110 to be maintained as constant or close to constant. This defines a condition noted C2 for the closed loop B2. Typically, means 11 for measuring this rotational speed are provided. Optionally, this condition C2 is replaced and/or complemented (depending on the steady regime considered) by a condition C2′ associated with the chosen steady regime and with a first monotonic and bijective function of the pitch setting of the vanes of the straightener 121 so as to impose a certain rotational speed of the propeller 110.

[0083] Very preferably, the method provides for a closed loop B1 to also regulate the flow rate of fuel injected into the combustion chamber 160 on the N2R/N1R ratio of the reduced rotational speeds of the high-pressure 130 and low-pressure 120 compressors. This closed loop B1 is a function of the steady regime selected on the basis of the appropriate device 2 in the aircraft turbomachine, in the sense that, preferably, this regime is previously associated with such a N2R/N1R ratio to be maintained as constant or close to constant. This defines a condition noted C1 regulating the closed loop B1. Typically, means 12, 13 for measuring the rotational speeds N1, N2, and the temperatures T2, T25 at the inlet of the low and high-pressure compressors 120 and 130 respectively are provided. Optionally, the condition C1 is replaced and/or supplemented (depending on the steady regime considered) by a condition C1′ associated with the chosen steady regime and with a second monotonic and bijective function of the pitch setting of the vanes of the straightener 121 so as to impose a certain N2R/N1R ratio.

[0084] The above-mentioned closed loop B1 is the preferred step (iii) of the method according to the invention. It allows the operating point of the low-pressure compressor 120 to be stabilised, or at least controlled, at the best compromise of efficiency and stability. Indeed, as illustrated in FIGS. 3 and 4 respectively, the steps (ii) and (i) of the method are likely to lead to a sliding of this operating point either towards low efficiencies or towards a certain instability. More specifically, each of these figures illustrates the operating field 99 of the low-pressure compressor 120, equipped with a reference frame of axis 97 and 98 corresponding respectively to the (optionally reduced) flow rate and the compression ratio of the low-pressure compressor 120. The pumping line 71 defines and/or estimates a stability limit of the latter. A margin (referred to as pumping margin) must be maintained between this pumping line 71 and the operating point of the low-pressure compressor 120, and in particular taken into account the uncertainties in measuring the operating parameters of the latter. A nominal and/or desired operating point is at the intersection of an iso-speed curve 90 and an operating line 80 corresponding to a flow rate at the inlet of the high-pressure compressor 130 (i.e. essentially a flow rate at the outlet of the low-pressure compressor 120). This operating point is located in an area 85 of optimum efficiency of the low-pressure compressor 120, at a sufficient margin from the pumping line 71.

[0085] FIG. 3 illustrates the impact that a variation in the pitch setting of blades of the propeller 110 is likely to have on the operating point of the low-pressure compressor 120 at the same rotational speed of the propeller 110. Such a variation changes the power requirement on the low-pressure shaft 101 of the aircraft turbomachine 100, and consequently the rotational speed of the high-pressure compressor 130. On the one hand, if the pitch setting of the blades of the propeller 110 opens from an initial position, the high-pressure compressor 130 is accelerated, resulting in a translation in the operating line of the low-pressure compressor 120 towards the line 82. The low-pressure compressor 120 then evolves towards lower efficiencies. On the other hand, if the pitch setting of the blades of the propeller 110 closes from this initial position, the high-pressure compressor 130 is slowed down, resulting in a translation in the operating line of the low-pressure compressor 120 towards the line 81. The low-pressure compressor 120 then evolves towards its stability limit.

[0086] FIG. 4 illustrates the impact of varying the pitch setting of the vanes of the straightener 121 on the operating point of the low-pressure compressor 120. In the same way as above, this variation affects the rotational speed of the high-pressure compressor 130. On the one hand, if the pitch setting of the vanes of the straightener 121 opens from an initial position so as to increase the air flow rate, the high-pressure compressor 130 is slowed down, which translates the operating point to the intersection of the operating line 81′ and the iso-speed curve 91 (which corresponds to the same speed as the curve 90 for the above-mentioned pitch setting of the vanes of the straightener 121). The low-pressure compressor 120 then evolves towards its stability limit. On the other hand, if pitch setting of the vanes of the straightener 121 closes from this initial position so as to decrease the air flow rate, the high-pressure compressor 130 is accelerated, which translates the operating point to the intersection of the operating line 82′ and the iso-speed curve 92 (which corresponds to the same speed as the curve 90 for the above-mentioned pitch setting of the vanes of the straightener 121). The low-pressure compressor 120 then evolves towards lower efficiencies.

[0087] In summary, the present invention relates to a method and a system for regulating the thrust of a high bypass ratio aircraft turbomachine 100, preferably at steady regime, by a direct action on a vane variable pitch setting system 123 for varying the pitch of vanes of a straightener 121 of a low-pressure compressor 120.

[0088] The present invention has been described above in connection with specific embodiments, which are illustrative and should not be considered limiting. In general, it will be apparent to a person skilled in the art that the present invention is not limited to the examples illustrated and/or described above.