GB2461718A - An aircraft propulsion arrangement having a single fan located within a curved guide path - Google Patents

Abstract

A propulsion arrangement of an aircraft comprises a single fan 3 in a guide path for propulsive flow, the guide path 1 having means to vary the propulsive flow rate in circumferential zones of the propulsive flow. Thereby, uniform total inlet pressure, balanced fan loading and equalized uniform exit velocity can be achieved. Ideally the means to vary the propulsive flow rate comprises a curved circular duct with the fan located at or near the axial position of maximum curvature. Alternatively or additionally, stator vanes 6, 7, and splitters may be arranged to reduce the fan inlet area for high velocity air and increase the inlet area for low velocity air flow. Preferably this arrangement is used at the rear of a blended wing body aircraft wherein boundary layer / wake air is ingested by the engines. The figure shows a side-view, with a further duct arrangement 2 behind duct 1, the guide ducts straddling elevator 5.

Description

AIRCRAFT PROPULSION ARRANGEMENT

The present invention relates to aircraft propulsion arrangements and more particularly to improving the efficiency of such propulsion.

Operation of turbine engines in general terms is relatively well known. Furthermore, it will be appreciated that there are continuing pressures to improve efficiencies and operational performance, both to provide a commercial advantage as well as greater environmental acceptability.

It is known that aircraft propulsion efficiency can be improved by ingesting boundary layer air into a fan or propeller. To propel an aircraft at constant speed, the engine thrust must be equal to the aircraft drag. In such circumstances, fuel is used dependent upon the amount of power required to drive the propulsor e.g. a fan or propeller, for propulsion. The power to drive a propulsor, a fan or propeller can be written as Integral (TxdSxdm + (UjxUj-UoxUo) /2xdm) In the above circumstances, useful power is derived by the expression thrust multiplied by aircraft speed, which in turn is given the relation Integral ( (Uj-Uo) xdmxUe) Where T is the exhaust temperature dS is the change of entropy in the propulsion system dm is the local mass flow Uj the local jet velocity Uo the inlet velocity to a propulsor PJ and P0 are assumed equal Ue is the aircraft speed In the above circumstances, it can be shown that if dS (change of entropy) is zero and the propeller or fan swallows the aircraft wake, the power to propel the aircraft would be as much as 20% less than a conventional propeller or fan. The propulsor is consequently provided with a non uniform inlet total pressure.

Although the above is known, it will be appreciated that providing a practical embodiment is difficult. The difficulty is primarily preventing the fan stalling or choking with the non uniform inlet total pressure.

In accordance with the present invention there is provided a propulsion arrangement of an aircraft, the arrangement comprising a fan in a guide path for propulsive flow, the guide path including means to vary propulsive flow rate in circumferential zones of the propulsive flow to balance total pressure loss in the ingested boundary layer.

Generally, the means to vary propulsive flow rate comprises a curved circular duct with the fan at or near the axial position of maximum curvature. Additionally, generally the maximum curvature is in the order of the radius of the duct. Typically, the curvature is in the order of 45°.

Normally, the guide path includes an inlet portion and an outlet portion with the fan between them. Generally, the fan is presented at a 45° angle to the guide path.

Possibly, the inlet portion and the outlet portion make a rectangular cross-section and the fan between them is located in a circular cross-section portion of the guide path.

Alternatively, the means to vary propulsive flow rate comprises circumferential splitters arranged in the guide path to make the fan inlet area smaller for any low velocity flow circumferential zones and larger for any high velocity flow circumferential zones in the axial air flow velocity generated by the single fan in order to balance the flow velocity across the propulsive arrangement.

Typically, the ducts are associated with the inlet guide vanes (IGV) and/or outlet guide vanes (OGV) respectively. These vanes control the work radially and circumfereritially to produce near constant exit velocity.

Also in accordance with the invention there is provided an aircraft propulsion assembly comprising a plurality of propulsion arrangements as described above.

Generally, each guide path will comprise a duct extending from an inlet to an outlet exhaust with the fan located in between. Normally, the inlets of the assembly will be interleaved with each other across an inlet front for the assembly. Typically, the exhausts for the ducts will be interleaved along a trailing edge of the propulsion assembly.

Embodiments of the present invention will now be described by way of example and with reference to the accompanying drawings in which: Fig. 1 is a schematic plan cross-sectional view of a first embodiment of a propulsive arrangement in accordance with the present invention; Fig. 2 is a schematic plan view of the propulsive arrangement depicted in Fig. 1; Fig. 3 is a schematic illustration of a configurational position for a propulsive arrangement in accordance with the present invention associated with an aircraft form; Fig. 4 is a schematic illustration of an exhaust end of the propulsive arrangement depicted in Fig. 1; and Fig. 5 is a schematic illustration of an inlet end of the propulsive arrangement depicted in Fig. 1.

It can be shown that as optimum design for a fan propulsive arrangement increases the velocity of every stream tube in the wake up to the same velocity. This velocity is not the aircraft speed as the drag of an aircraft includes lift induced drag which is not seen as a wake. It can further be shown that the loss of performance if the jet is not exactly constant is small. The optimum is relatively flat. If the wake is mixed before the fan much of the inherent gain is lost due to mixing loss and more if some form of Stirrer' is used. In principle, the propeller or fan would have to increase the velocity of the flow at the wake where the velocity was lowest by the largest amount. It can be shown by conventional aerodynamics that this is impossible to achieve efficiently unless the propulsor had many stages at the bottom of the wake and one at the edge. This is necessary to keep the pressure rise to a reasonable proportion of the relative inlet dynamic head. The current invention requires only a single stage fan to satisfy the variable work through the boundary layer to produce a reasonable flat exit profile with inlet velocities going to zero at the centre of the wake. Given a typical flight mach number of O.85M and assuming the induced drag is equal to the boundary layer drag the highest pressure ratio would be about 1.9 and the lowest about 1.25. To produce the former, a supersonic inlet mach number fan would be needed as a single stage.

It is well known that supersonic inlet flow fans have efficiencies that fall more or less linearly with increasing incidence and have a falling efficiency with falling pressure at constant incidence when operating choked or at unique incidence.

The upstream wake can pass through a slotted inlet along the aircraft trailing edge before transfer to a series of circular ducts. The wake profile can be mapped onto a plane where the fans sit. At the fan inlet there is an inlet guide vane, at any radius there is a circumferential variation in exit angle to provide the substantially constant nozzle exit velocity. The angle of the inlet guide vanes being set so that at every point the relationship DH/Ur/Ur=l_Va/Ur*(tan(A2)+tan(Ao)) holds; Where Ur is the rotor speed at that radius; DH the local enthalpy rise required to produce the substantially constant nozzle exit velocity Ur2; A2 is the rotor exit angle assumed near constant circumferentially; Ao is the IGV exit angle.

Thus, given that DH, Va and Ur for a fan design are conventional and Ao is prescribed once the mean level is chosen it can be stated The inlet angle to a rotor Al=atan(DH/Ur/Va+tan(A2)) The rotor loss which varies nearly linearly with Al-Al (choke) is then clearly excessive. Thus as Al max DHmax/Ur/Vamin-DHmin/Ur/Vamax then it will be understood that the fundamental requirement is that DHmax does not coincide with Vamin and vice versa.

In the above circumstances, the requirement for the ideal optimum efficiency boundary layer propulsor leads apparently to a necessity for a poor fan or propeller for the present propulsion arrangement.

Use by the present propulsion arrangement of DH variation gives most of the ideal gain but limits the maximum and minimum values. However, there is a need to modify the axial velocity to reduce the range and in particular we need to increase the velocity at the bottom of the wake and reduce it at the edge since the mach number at the edge is considered too high and at the bottom too low. This can be done in two ways.

Firstly, if the wake is allowed to flow into a curved circular duct with a fan at or near the axial position of maximum curvature and if the sharpest curvature is of the order of the duct radius, it can be shown that the axial velocity at an inlet to a fan is essentially uniform even with a total pressure gradient corresponding to a wake.

Furthermore, the ideal total pressure rise across the fan which is equal to a constant exit dynamic head minus the local dynamic head in the wake is directly proportional to the static pressure change in the inlet duct and the flow downstream of the fan is constant. The bend accelerates the low velocity flow at the bottom of the wake and diffuses the high mach number flow at the edge. Downstream of the fan a further opposite bend is needed to adjust or compensate for the thrust line variations.

Alternatively, or additionally, IGV5 and OGV5 are fitted with part circumferential splitters, so a similar effect can be achieved by making the fan inlet area smaller for the low velocity upstream flow and larger for the high velocity upstream flow. This causes less flow in the high velocity region and more in the low velocity region with the result that the static pressure at exit is constant to satisfy the Kutta condition.

In order to optimise the arrangement there are three variables that can be used.

a) Mean Whirl Angles In general, the highest mach number is also the maximum positive incidence. To reduce the loss increasing the positive whirl at the point on the circumference where the mach number is lowest also reduces the mach number where it is highest. it is possible to reduce the lowest mach number to subsonic where the loss curves become more rounded and the lowest mach number points can be run at negative incidence.

b) Duct Curvature Duct curvature can be increased beyond the uniform velocity value and the mach number can be higher effectively in the centre of the wake than the edge flow when in the duct.

c) Non Optimum Exit Velocities The effect of non optimum exit velocities can be readily accounted for so it can be used to cptimise the whole system by individually treating each portion of the wake in a near ideal fashion has a benefit of lower fuel burn than a conventional engine taking in free stream air by curving the duct. By optimising the IGV and rotor blade speed it is possible to design a fan where every element operates with a DH/Ur/Ur and Va/Ur in the normal safe operating region for fans without stall. A single stage fan operates with an incidence variation around the circumference and this means more loss.

The use of a fan inlet shock model and an IGV and stator loss model predicts that this fan will be less efficient than one designed to do the average or mean duty.

Thus, there is a benefit of lower fuel burn over a conventional propulsion system but a known fan efficiency.

The present propulsion arrangement is best suited to an aircraft with Delta or blended wing body.

A blended wing body claims a drag advantage over conventional winged aircraft so such a combination with a present boundary layer propulsion arrangement could approach the Acara target.

Generally, the upper and lower wakes need to be treated in a similar but different way the lower wake needs to be bent in the opposite sense.

Normally, the upper surface boundary layer is of the order of a metre thick whilst the lower surface has less of a boundary layer so it is envisaged that there will be a significant number of fans in present propulsion arrangements along a wing trailing edge in a practical use of the present propulsive arrangements. This is likely to lead to less drag than conventional podded engines.

It will be appreciated from the above that the present invention relates to propulsive inefficiencies with regard to fan operation in terms of the different propulsive air flow rates generated by different parts of the driving fan blades. It will be understood that the air flow rate at the centre of the rotating fan blade assembly will generally be much lower than that at the tips of the fan blades leading to problems with respect to boundary layer separation with previous single fan propulsive arrangements. The present propulsive arrangements attempt to equalise the flow rates across the single fan through utilisation of curvature in a duct and/or through use of flow splitter devices associated generally with the inlet guide vanes or outlet guide vanes. In short, by utilisation of these ducts or splitters in the air flow guide paths, it is submitted that practical propulsive air flow rates through the propulsive arrangement can be adjusted for consistency across the width of the propulsive arrangement.

Typically, a practical propulsive arrangement in accordance with the present invention will combine both the duct and the flow splitter devices in order to create the necessary equalisatiori of propulsive air flow across the fan.

Fig. 1 illustrates schematically a propulsive arrangement in accordance with the present invention. By considering the air flow A through the curved ducts 1, 2 with respective single fans 3 in each duct 1, 2 in terms of circumferential flow zones Aa, Ab, Ac in each duct 1, 2.

It will be noted that the inherent differential in air flow rate between different parts of the fan 3 in each duct 1, 2 is arranged through the bending curvature of the ducts 1, 2 such that the outer flow Aa becomes an inner flow AAa either side of an exhaust trailing edge of the wing and an inner surface flow Ac in each duct 1, 2 becomes an outer circumferential zone flow AAc at the exit end of the propulsive arrangement. Each of these circumferential flow zones Aa, Ac is incident upon the high speed peripheral regions of the fan 3, whilst central zones Ab are incident upon the slower rotating fan blades of the fan 3 and constitute the central parts of the output flow AA from the propulsive arrangement. However, due to the curved nature of the ducts 1, 2 and appropriate shaping of the ducts 1, 2, it will be appreciated that these flows Ab travel less distance and therefore provide flow rates which are equalised with the flow rates AAa, AAc at the outlet side of the propulsive arrangement.

It will be noted that there is a 45 degree orientation of each fan 3 to the exhaust duct 12.

The ducts 1, 2 are designed such that as depicted in Fig. 2, in a first intake portion bA, bOB of the ducts 1, 2, a transition from a rectangular intake cross section to a circular cross section fan portion hA, biB is provided in which the fan 3 (Fig. 1) operates in order to generate propulsive air flow. Subsequent to the fan portions hA, biB, an exhaust portion 12A, 12 B is provided in which the circular duct flow is translated to a rectangular propulsive flow. In such circumstances within each duct 1, 2 through the bend in those ducts 1, 2 there is equalisation of propulsive flow rate despite variations in the propulsive effect of the respective fans 3 upon different parts of the incident flow A. It is critical that the single fan 3 is appropriately positioned as illustrated in Fig. 1 and described above at the position of maximum curvature and typically at 45° to the general axis of presentation of internal air flow A. In such circumstances, the blades of the fan 3 are then appropriately presented to different portions Aa, Ab, Ac to generate the necessary propulsive flows in accordance with the present invention.

In addition to, or in some instances, as an alternative to providing ducts 1, 2 with appropriate curvature, it is possible to provide flow splitters in the respective guide paths provided for the propulsive flow.

Thus, as also illustrated in Fig. 1, a flow splitter may be incorporated into inlet guide vanes 6 for the fan 3.

Whilst outlet guide vanes 7 for the fan 3 will also be provided as flow splitters. In such circumstances, these flow splitters will differentially restrict air flow in differing circumferential zones generated by the fan 3, such that as described above, essentially the potential fan inlet area is smaller for low velocity circumferential zones (Ab) in comparison with higher velocity flows (Aa, Ac) such that there is less flow in the high velocity regions and more flow in the low velocity regions than would naturally occur in order to create a substantially static pressure at the exit as defined above.

Fig. 3 provides a schematic plan view illustrating position of propulsion arrangements in accordance with the present invention upon a representative delta wing aircraft 20. Thus, as can be seen, the propulsive arrangements 21 are generally located along a rear surface edge of the representative aircraft form 20. As indicated previously, this should reduce drag associated with providing propulsive arrangements in pods upon pylons below aircraft wings utilised previously.

The propulsion arrangements 21 act as propulsors which are interleaved at the trailing edge of the wing of the aircraft 20 which as indicated will generally be a blended or Delta wing aeroplane. Generally, the propulsors are arranged along the trailing edge of the wing such that their intakes are provided alternatively above and below the wing whilst their respective exhausts similarly alternate above and below the wing of the aircraft 20. The propulsor ducts pass through the wing such that fan portions in the duct of the propulsors are generally circular in cross section and arranged at an angle to intake portions and exhaust portions of the propulsor duct path.

Figures 4 and 5 respectively show exhaust configuration and inlet configurations for a wing in accordance with the current invention.

As can be seen in Fig. 4 the duct paths 1, 2 are arranged in an interleaved relationship along the trailing edge 5 of the wing. Similarly in Fig. 5 it can be seen that the inlets 4 for the ducts which lead to the exhaust portions 12 are interleaved to provide appropriate load distribution from the trailing edge 5.

By interleaving the inlet portions 10 and outlet portions 12 it will be understood that in addition to providing equalisation along the propulsor bank of the wing at the trailing edge 5 it may also be possible to more readily accommodate these guide paths formed by the portions 10, 11 with a fan section between them within the profile of a wing.

As illustrated in Fig. 2 generally a large number of propulsion arrangements and propulsors will be provided on the trailing edge of the blended wing aircraft. The number of propulsors will depend upon performance requirements but as illustrated in Figures 4 and 5 these propulsors will be intermingled and interleaved to create the desired flow bypass for propulsion.

Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon.

Claims (16)

1. CLAIMS1. A propulsion arrangement of an aircraft, the arrangement comprising a single fan in a guide path for propulsive flow, the guide path including means to vary propulsive flow rate in circumferential zones of the propulsive flow to balance the fan loading due to the inlet total pressure variation.
2. 2. An arrangement as claimed in claim 1 wherein the means to vary propulsive flow rate comprises a curved circular duct with the fan at or near the axial position of maximum curvature for a near constant exit velocity
3. 3. An arrangement as claimed in claim 2 wherein the maximum curvature is in the order of the radius of the duct.
4. 4. An arrangement as claimed in claim 2 or claim 3 wherein the curvature is in the order of 45°.
5. 5. An arrangement as claimed in the preceding claim where the guide path includes an inlet portion and an outlet portion with the fan between them.
6. 6. An arrangement as claimed in claim 5 where the inlet portion and the outlet portion make a rectangular cross-section and the fan between them is located in a circular cross-section portion of the guide path.
7. 7. An arrangement as claimed in any preceding claim where the fan is presented at a 45° angle to the guide path.
8. 8. An arrangement as claimed in any preceding claim wherein the means to vary propulsive flow rate comprises circumferential splitters arranged in the guide path to make the fan inlet area smaller for any low velocity flow circumferential zones generated by the single fan and larger for any high velocity flow circumferential zones generated by the single fan in the axial air flow velocity generated by the single fan in order to balance the flow velocity across the propulsive arrangement.
9. 9. An arrangement as claimed in claim 8 wherein the splitters are associated with the inlet guide vanes (IGV) and/or outlet guide vanes (OGV) respectively.
10. 10. A propulsion arrangement for an aircraft substantially as hereiribefore described with reference to the accompanying drawings.
11. 11. An aircraft propulsion assembly comprising a plurality of propulsion arrangements as claimed in any preceding claim.
12. 12. An assembly as claimed in claim 11 herein each guide path will comprise a duct extending from an inlet to an outlet exhaust with the fan located in between.
13. 13. An assembly as claimed in claim 11 or claim 12 wherein the inlets of the assembly will be interleaved with each other across an inlet front for the assembly.
14. 14. An assembly as claimed in any claims 11 -13 wherein the exhaust for the ducts will be interleaved along a trailing edge of the propulsion assembly.
15. 15. An aircraft including a propulsion arrangement as claimed in any preceding claim.
16. 16. Any novel subject matter or combination including novel subject matter disclosed herein, whether or not within the scope of or relating to the same invention as any of the preceding claims.