GB2542184A

GB2542184A - Aircraft comprising a boundary layer ingesting propulsor

Info

Publication number: GB2542184A
Application number: GB1516106.0A
Authority: GB
Inventors: Popovic Ivan; Ful Hen Chong Ellis; G Menon Kalyani; j armstrong Michael
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2015-09-11
Filing date: 2015-09-11
Publication date: 2017-03-15
Also published as: GB201516106D0

Abstract

An aircraft 10 comprises a propulsion arrangement 50 comprising first and second propulsors 2 and 3. The first propulsor 2 comprises an inlet (29, Fig 2) and an exhaust (35, Fig 2) upstream of an inlet (37, Fig 2) of the second propulsor 3 such that an exhaust of the first propulsor is ingested by the second propulsor. The first propulsor has an inlet area less than an inlet area of the second propulsor and is arranged to ingest at least part of a boundary layer flow adjacent a surface 13 of the aircraft.

Description

Aircraft Comprising a Boundary Layer Ingesting Propulsor

The present disclosure concerns an aircraft, particularly an aircraft having boundary layer ingesting propulsors.

It is known to increase propulsive efficiency of an aircraft providing propulsors (such as ducted fans or propellers) which ingest boundary layer air. Boundary layer air extends from a wetted surface to a thickness normally defined as the distance from the wetted surface at which the viscous flow velocity is 99% of the freestream velocity (the surface velocity of an inviscid flow). Consequently, boundary layer air moves more slowly than the freestream flow. As such, the propulsors will accelerate the airflow to a greater extent for the same exhaust velocity, thereby increasing propulsive efficiency. Where such boundary layer propulsors are provided at an aft end of the aircraft, such propulsors also provide “wake filling”, thereby decreasing aircraft drag.

Aircraft have been proposed which take advantage of these effects by placing one or more propulsors at an aft end of the aircraft. A first example includes the Sax40 proposed by the Silent Aircraft Initiative and described in Airframe Design for "Silent Aircraft", J Hileman, Z Spakovszky, M Drela and M Sargeant, AIAA-2007-0453, 45th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada. A second example includes the NASA D8, described in External Aerodynamics Simulations for the MIT D8 “Double-Bubble” Aircraft Design, S Pandya, Seventh International Conference on Computational Fluid Dynamics (ICCFD7), Big Island, Flawaii, July 9-13, 2012.

Flowever, the aerodynamic benefits of such aircraft are relatively small. Furthermore, several previous proposals place the propulsor inlet spaced from a single surface, (such as the top of the fuselage in the D8), resulting in inlet flow distortion. Such distortion may cause high levels of vibration, and large loads on the fan. Consequently, such an arrangement will require a distortion tolerant fan, which presents further design challenges. Such arrangements may also result in a relatively noisy propulsion system, which may not be acceptable to regulators, passengers, or the public. Similar considerations apply where the boundary layer ingesting propulsor is provided at the rear of a conventional tubular fuselage, in view of the higher velocity inlet air at the rotor tips compared to at the roots.

The term “longitudinal” used herein refers to an axis of the aircraft extending from a nose of the aircraft to a tail of the aircraft. The term “lateral” refers to an axis running perpendicular to the longitudinal axis in the horizontal plane when the aircraft is on the ground or in normal level flight. The term “vertical” refers to an axis relative to when the aircraft is on the ground or in normal level flight.

According to the present disclosure there is provided an aircraft comprising a propulsion arrangement comprising first and second propulsors, wherein the first propulsor comprises an inlet and an exhaust upstream of an inlet of the second propulsor such that an exhaust of the first propulsor is ingested by the second propulsor, the first propulsor having an inlet area less than an inlet area of the second propulsor and being arranged to ingest at least part of a boundary layer flow adjacent a surface of the aircraft.

Advantageously, the second propulsor ingests substantially only freestream air and the exhaust from the first propulsor. Consequently, the boundary layer is accelerated (thereby providing the advantages outlined above), while precluding the requirement for a large, distortion tolerant propulsor, as the majority of the propulsive thrust can be produced by the downstream, none boundary layer ingesting propulsor.

The propulsion arrangement may be mounted such that the inlet of the first propulsor is located at a downstream end of a fuselage of the aircraft.

The propulsion system may comprise a controller configured to control operation of one or both of the first and second propulsors. The controller being configured to control the first propulsor to provide exhaust fluid flow such that fluid flow at the second propulsor inlet has the same or a similar velocity to freestream air. Consequently, the velocity profile of air entering the second propulsor is substantially constant across the radius of the second propulsor inlet.

The propulsion system may comprise a driving arrangement configured to drive the first and second propulsors. The drive arrangement may comprise an internal combustion engine such as a gas turbine engine. The internal combustion engine may be mechanically coupled to the second propulsor to mechanically drive the second propulsor. The second propulsor and internal combustion engine may be coupled via a reduction gearbox.

The first propulsor may be driven by an electric motor, which may be powered by electrical power provided by a generator driven by the internal combustion engine.

The first and second propulsors may be configured to counter-rotate relative to one another.

One or both of the first and second propulsors may comprise a bladed rotor. The propulsion arrangement may comprise a duct configured to enclose a radially outer end of at least one of the first and second propulsors.

The propulsion arrangement may comprise an inlet guide vane located upstream of the second propulsor. The inlet guide vane may comprise a variable angle inlet guide vane configured to vary the angle of incidence of air flow ingested by the second propulsor. The inlet guide vane may extend radially inwardly from the duct, and may extend coaxially with the bladed rotor of the first propulsor such that a radially inward end of the inlet guide vane encloses a radially outer end of the bladed rotor of the first propulsor. Advantageously, the combination of the inlet guide vane and first propulsor can provide consistent inlet flow for the second propulsor, accelerating boundary layer flow while removing swirl and preventing end vortices rom the first propulsor bladed rotors from forming.

The propulsion arrangement may comprise an outlet guide vane located downstream of the second propulsor. Advantageously, swirl produced by the propulsion arrangement can be reduced or removed, thereby increasing propulsive efficiency.

The skilled person will appreciate that except where mutually exclusive, a feature described in relation to any one of the above aspects of the invention may be applied mutatis mutandis to any other aspect of the invention.

Embodiments of the invention will now be described by way of example only, with reference to the Figures, in which:

Figure 1 is a plan view of an aircraft including a first propulsion arrangement in accordance with the present disclosure;

Figure 2 is a cross sectional view through the area surrounded by dotted lines in figure 1;

Figure 3 is a plan view of a second propulsion arrangement in accordance with the present disclosure; and

Figure 4 is a cross sectional view through the area surrounded by dotted lines in figure 3.

With reference to Figure 1, an aircraft 10 in accordance with the present disclosure comprises a fuselage 12, a pair of wings 14, an empennage including vertical 15 and horizontal 17 tail surfaces and a propulsion arrangement 50. The fuselage 12 defines a nose 16 at a forward end, and a tail 18 at an aft end. A notional line extending horizontally between the nose 16 and tail 18 defines a longitudinal centre line X. A horizontally extending (when the aircraft is on the ground or in level flight) line extending perpendicular to the longitudinal centre line defines a lateral axis Y. A pair of main engines 19 are provided in the form of gas turbine engines mounted on each wing 14. A rear portion of the fuselage 12 includes an inwardly tapering portion 13. The inwardly tapering portion 13 transitions between a generally cylindrical substantially constant cross section portion 11 and the tail 18, which meets the longitudinal axis X.

As shown in figures 1 and 2, a boundary layer (represented by diagonal hatching 5) is defined by a region adjacent an outer surface of the aircraft 10, such as fuselage 12. The boundary layer has a thickness defined as as the distance from the wetted surface at which the viscous flow velocity is 99% of the freestream velocity. The boundary layer thickness δ at a given point on a wetted surface is dependent on an axial distance x from an upstream end (i.e. a leading edge) of the respective component, a freestream velocity U adjacent the wetted surface, and the viscosity v of the fluid approximately in accordance with the following equation:

Equation 1

In one example, for an aircraft (in this case, an Airbus A320) having a length of 37.5 metres, a cruising speed of 230 metres per second at an altitude of 11,000 metres, and a viscosity of air at this altitude under standard atmospheric conditions being 1.458 10"5 N.s/m2, the corresponding boundary layer thickness at the tail of fuselage at cruise is approximately 50 cm.

In some cases, the boundary layer thickness may be increased due to turbulent boundary layer separation, caused for example by high angles of attack of the surface relative to the relative wind. In any case, the boundary layer thickness of the aircraft in cruise can be easily ascertained for a given aircraft design.

Figure 2 illustrates the propulsion arrangement 50 in more detail. The propulsion arrangement 50 comprises first and second propulsors. Each propulsor comprises a bladed rotor 2, 3, which is configured to rotate about a respective rotational axis, (in this case the longitudinal axis X) to thereby generate thrust in use. Each bladed rotor comprises a plurality of blades 28 30 mounted to a hub 32. Each blade 28, 30 defines an inner radius at a root defined by an outer circumference of the hub 32 at the corresponding location, and an outer radius defined by a tip 34, 36. The hub 32 is generally inwardly tapered in a direction extending aft along the longitudinal axis, such that the inner radius of the blades 28 of the first rotor 2 is generally greater than the inner radius of the blades 30 of the second rotor 3.

The first bladed rotor 2 comprises a plurality of blades having a diameter such that the respective blade tip 34 extends radially from the fuselage a distance approximately equal to the boundary layer thickness δ, i.e. the distance from the root to the tip 34 is equal to the boundary layer thickness δ. In some cases, the first bladed rotor 2 extends a distance less than the boundary layer thickness δ, i.e. equal to a portion only of the boundary layer thickness. This is because the majority of the velocity deficit in the boundary layer is generally found close the fuselage surface, with greater than 50% of the velocity deficit being found in the 30% of the boundary layer closest to the fuselage. Consequently, at least some of the benefits of the disclosed arrangement can be provided by a first propulsor having a relatively small diameter rotor blade, and a relatively low power requirement. Consequently, in either case, an inlet 29 of the first propulsor 2 ingests boundary layer air, and accelerates this air to provide an exhaust flow at an exhaust 35. Each blade 28 may comprise a variable camber, pitch or twist along its span, such that the thrust produced by the blade varies along the span, generally decreasing in a direction from the root to the tip 34. Consequently, each blade 28 of the first propulsor 2 accelerates flow close to the fuselage 12 surface to greater extent than flow further from the fuselage surface, in proportion to the velocity deficit within the boundary layer at that point. Consequently, the second propulsor 3 receives inlet flow with relatively little radial variation in velocity.

The second bladed rotor 3 is located downstream of the first bladed rotor 2, having an inlet 37 located downstream of the exhaust 35 of the first rotor 2. The second bladed rotor 3 has an outer radius greater than the first bladed rotor 2, such that the blades 30 ingest air exhausted by the first bladed rotor 2, and freestream air. In general, the thrust produced by the second propulsor is therefore greater than that provided by the first propulsor. As can be seen, the air exhausted by the first bladed rotor 2 includes turbulent air 70 as a result of tip vortices shed by the tips 34 of the blade 28 of the first propulsor 2.

Consequently, the second propulsor 3 must be designed to accommodate these vortices.

The propulsion arrangement 50 further comprises an internal combustion engine in the form of a gas turbine engine 20. The gas turbine engine 20 comprises a compressor 22, which is configured to compress an inlet air flow from an inlet (not shown). A combustor 24 is provided downstream of the compressor 22. The combustor 24 receives compressed air from the compressor 22, mixes it with fuel, and combusts it to produce heated exhaust fluid. A turbine 26 is provided downstream of the combustor 24, and receives the heated exhaust fluid in use, using the exhaust fluid to turn the turbine 26. A first shaft 38 couples the turbine 26 to the compressor 24, such that rotation of the turbine 26 in use causes rotation of the compressor 24. The first shaft 38 is coupled to the second propulsor rotor 3. Consequently, the second propulsor is directly driven by the gas turbine engine 20.

The first shaft 38 is also coupled to an electrical generator 40, to thereby generate electrical power. The electrical generator is electrically coupled to an electric motor 42 via an electrical interconnector 44. An output shaft 46 of the electric motor 42 is coupled to the first propulsor rotor 2 to thereby drive the first propulsor 2. The output shaft 46 is arranged coaxially with the first shaft 38, such that the first shaft 38 is nested within the output shaft 46 of the gas turbine engine 20. In this way, the first propulsor 2 is driven by an electric motor using electrical power generated by mechanical power provided by the gas turbine engine 20, while the second propulsor 3 is driven directly by the gas turbine engine 20. Consequently, the impact of losses associated with electrical drive shafts is minimised, since the power requirement of the first propulsor 2 is significantly less than the power requirement of the second propulsor 3. The electrical generator 40 is also coupled to further aircraft electrical systems, such as avionics, pumps, control surface actuators etc. Consequently, the gas turbine engine 20 and generator 40 act as an Auxiliary Power Unit (APU), to thereby provide at least backup power in the event of failure of one of the main engine driven electrical generators.

Consequently, in view of the separate shafts 38, 46 and driving means driving the first and second propulsor rotors 2, 3, each rotor 2, 3 can be driven at a speed independent of the other. The speed of the second rotor 3 can be controlled by varying the rotational speed of the gas turbine engine 20, while the speed of the first rotor 2 can be varied by varying the speed of the electric motor 44. The first and second propulsors 2, 3 are configured to counter-rotate in use.

Each rotor 2, 3 further comprises a blade pitch actuator 48, configured to vary the pitch of the blades of the respective rotor 2, 3. Consequently, the thrust of the propulsors 2, 3 can be controlled independently of rotational speed and forward airspeed to some degree. In general, the first and second propulsor 2, 3 blade pitch will be controlled in order to provide maximum aerodynamic efficiency, while preventing excessive turning of the airflow by either propulsor (which may result in stalling of one or both of the propulsors 2, 3).

The propulsion arrangement 50 comprises a controller 72 configured to control the first and second propulsors 2, 3, including the gas turbine engine 20, electrical motor 42, and variable pitch actuators 28, 48.

In a first control method, the controller 72 controls the motor 42 and variable pitch actuators 28 to produce a first propulsor exhaust flow having a velocity across the radius of the exhaust substantially equal to the freestream velocity at the second propulsor inlet 37. Meanwhile, the controller 72 controls the gas turbine engine 20 to provide the required driving torque for the electrical generator 40 to supply the necessary electrical power to the motor 42, and to drive the second propulsor 3 at the necessary speed, such that the second propulsor 3 generates sufficient thrust in combination with the first propulsor 2 to match a thrust demand. The thrust demand may be varied in accordance with aircraft speed and flight profile. For example, the propulsion arrangement 50 may be utilised on takeoff to provide additional takeoff thrust, but may not be operated (or may be operated at low power) during climb. At cruise (i.e. at high speed), the propulsion arrangement 50 may be operated again, in order to provide the greatest benefit. The controller 72 controls the second pitch actuator 48 to maintain the second propulsor 3 at an efficient working line, as would be understood by the skilled person.

Figures 3 and 4 show an alternative propulsion arrangement 150. The arrangement 150 comprises first and second propulsors 102, 103, comprising bladed rotors similar to those of the propulsors 2, 3. Each propulsor is driven in a similar manner to the propulsion arrangement 150. However, the propulsion arrangement 150 differs from the arrangement 50 in that the arrangement 150 comprises a duct 152 provided radially outwardly of the first and second propulsors 102, 103, and shown more clearly in figure 4.

The duct 152 comprises an annular hollow tube. The duct 152 comprises an annularly outer wall having an outer surface 154 and an inner surface 156 joined by leading and trailing edges 158, 160 defining an annular aerofoil cross-sectional profile. The nacelle 152 is coaxial with the first and second propulsor 102, 103 rotational axes.

The duct 152 is connected to the tail 116 of the fuselage 112 by an outlet guide vane 162. The outlet guide vane 162 has an aerofoil cross sectional profile, and extends radially between the tail 116 and the inner surface 156 of the nacelle 152 downstream of the second propulsor 103. The outlet guide vane 162 serves to both support the nacelle 152, and reduce exit swirl produced by the propulsors 102, 103. The outlet guide vane 162 comprises an optional pitch actuator 164 configured to alter the pitch of the outlet guide vane 162 in accordance with varying flight conditions, propulsor rotational speeds, and propulsor pitch settings in order to substantially cancel propulsor exit swirl.

The duct further comprises an inlet guide vane 166. The inlet guide vane 166 has an aerofoil cross sectional profile, and is located coaxially in the same longitudinal plane as the first propulsor 102 rotor blades 128. The inlet guide vane 166 extends radially between inner surface 156 of the nacelle and a point spaced from a tip 134 if the first propulsor 102 rotor blades 128. The inlet guide vane 166 may comprise a radially inner annular wall extending around the circumference of the first propulsor 102. Again, the inlet guide vane 166 comprises a pitch actuator 168, configured to alter the pitch of the inlet guide vane 166.

The purpose of the inlet guide vane 166 is as follows. Firstly, pitch of the inlet guide vane 166 can be adjusted to alter the incoming relative angle of the freestream air approaching the second propulsor 103. Secondly, the inlet guide vane 166 prevents tip vortices from being shed by the first propulsor 102. If the inlet guide vane 166 were not present (as demonstrated by line 70 in figure 2), then tip vortices 170 would be shed from the first propulsor 2, resulting in turbulent entry flow to the second propulsor 103. Advantageously, the duct 152 and inlet guide vane 166 reduces distortion to the second propulsor 103 still further compared to the arrangement of the first embodiment.

Accordingly, the present disclosure describes a propulsion system which provides the high propulsion efficiency of a boundary layer ingesting propulsor, while not requiring a distortion tolerant propulsor rated for the full thrust requirement of the airframe. Furthermore, since the thrust is shared between two propulsors, the diameter of the propulsors can be reduced for a given thrust. Consequently, the risk of a tail strike may be reduced.

It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and subcombinations of one or more features described herein.

For example, the or each propulsor may be directly driven via a reduction gearbox. The variable pitch actuators could be omitted. The first and second propulsors could be driven using different arrangements, such as either or both being driven mechanically and / or by an electric motor, and could be mechanically coupled to be driven at the same velocity.

Claims

1. An aircraft (10) comprising a propulsion arrangement (50) comprising first and second propulsors (2, 3), wherein the first propulsor (2) comprises an inlet (29) and an exhaust (35) upstream of an inlet (37) of the second propulsor (3) such that an exhaust of the first propulsor (2) is ingested by the second propulsor (3), the first propulsor (2) having an inlet area less than an inlet area of the second propulsor (3) and being arranged to ingest at least part of a boundary layer flow adjacent a surface (13) of the aircraft (10).

2. An aircraft according to claim 1, wherein the propulsion arrangement is mounted such that the inlet (29) of the first propulsor (2) is located at a downstream end (13) of a fuselage (12) of the aircraft (10).

3. An aircraft according to claim 1 or claim 2, wherein the propulsion system (50) comprises a controller (72) configured to control operation of one or both of the first and second propulsors (2,3), the controller (72) being configured to control the first propulsor (2) to provide exhaust fluid flow such that fluid flow at the second propulsor inlet (35) has the same or a similar velocity to freestream air.

4. An aircraft according to any of the preceding claims, wherein the propulsion system comprises a driving arrangement configured to drive the first and second propulsors (2,3), the drive arrangement comprising an internal combustion engine such as a gas turbine engine (20).

5. An aircraft according to claim 4, wherein the internal combustion engine (20) is mechanically coupled to the second propulsor (3) to mechanically drive the second propulsor (3).

6. An aircraft according to any of the preceding claims, wherein the first propulsor (2) is driven by an electric motor 42.

7. An aircraft according to claim 6 when dependent on claim 5, wherein the electric motor (42) is powered by electrical power provided by a generator (40) driven by the internal combustion engine (20).

8. An aircraft according to any of the preceding claims, wherein the first and second propulsors (2, 3) are configured to counter-rotate relative to one another.

9. An aircraft according to any of the preceding claims, wherein the propulsion arrangement (150) comprises a duct (152) configured to enclose a radially outer end of at least one of the first and second propulsors (102, 103).

10. An aircraft according to any of the preceding claims, wherein the propulsion arrangement (50) comprises an inlet guide vane (166) located upstream of the second propulsor (103), wherein the inlet guide vane (166) may comprise a variable angle inlet guide vane configured to vary the angle of incidence of air flow ingested by the second propulsor (103).

11. An aircraft according claim 10 when dependent on claim 9, wherein the inlet guide vane (166) extends radially inwardly from the duct (152), and may extend coaxially with a bladed rotor (134) of the first propulsor (102) such that a radially inward end of the inlet guide vane (166) encloses a radially outer end of the bladed rotor (134) of the first propulsor (102).

12. An aircraft according to any of the preceding claims, wherein the propulsion arrangement (150) comprises an outlet guide vane (162) located downstream of the second propulsor (103).