EP2372109A2

EP2372109A2 - Diffusive nacelle

Info

Publication number: EP2372109A2
Application number: EP11152869A
Authority: EP
Inventors: Erminio Zanenga; Roland Schweikhard
Original assignee: Rolls Royce Deutschland Ltd and Co KG; Rolls Royce PLC
Current assignee: Rolls Royce Deutschland Ltd and Co KG; Rolls Royce PLC
Priority date: 2010-02-09
Filing date: 2011-02-01
Publication date: 2011-10-05
Also published as: US20110194932A1; GB201002059D0; GB201002307D0; EP2372109A3

Abstract

A nacelle for a propeller gas turbine engine, the nacelle comprising a forebody upstream of the propellers and an afterbody downstream of the forebody. The forebody comprises a first, upstream region and a second, downstream region. The first region has a convex profile and the second region has a concave profile to cause diffusion of the airflow into the propellers.

Description

The present invention relates to a diffusive nacelle for a propeller gas turbine engine.
Referring to Figure 1, a conventional twin-spooled, contra-rotating propeller gas turbine engine is generally indicated at 10 and has a principal rotational axis 9. The engine 10 comprises a core engine 11 having, in axial flow series, an air intake 12, an intermediate pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, an intermediate pressure turbine 18, a free power (or low-pressure) turbine 19 and a core exhaust nozzle 20. A nacelle 21 generally surrounds the core engine 11 and defines the intake 12 and nozzle 20 and a core exhaust duct 22. The engine 10 also comprises two contra-rotating propeller stages 23, 24 attached to and driven by the free power turbine 19 via shaft 26. The configuration having the propeller stages 23, 24 towards the rear of the gas turbine engine 10 is termed a "pusher" configuration, as opposed to the "puller" or "tractor" configuration having the propeller stages 23, 24 towards the front of the engine 10.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 12 is accelerated and compressed by the intermediate pressure compressor 14 and directed into the high-pressure compressor 15 where further compression takes place. The compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high-pressure, intermediate pressure and free power turbines 17, 18, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high-pressure, intermediate pressure and free power turbines 17, 18, 19 respectively drive the high and intermediate pressure compressors 15, 14 and the propellers 23, 24 by suitable interconnecting shafts. The propellers 23, 24 normally provide the majority of the propulsive thrust. In the embodiments herein described the propellers 23, 24 rotate in opposite senses so that one rotates clockwise and the other anti-clockwise around the engine's rotational axis 9.
One problem with a conventional pusher propeller gas turbine engine 10 is that its cruise speed is limited to slightly below transonic, predominantly due to the drag rise encountered when flying at higher speeds. One of the main causes of this drag rise is that generally the root of each blade forming the propeller stages 23, 24 can not be shaped with the thin profiles required for high speed. The root has to be thick enough to guarantee the structural robustness of the blades given the high aerodynamic and mechanical loads acting on the propeller stages 23, 24, which disadvantageously adds significant weight to the engine 10. The airflow passing between the blade roots may easily become supersonic if the propeller gas turbine engine 10 operates at transonic cruise speed, around Mach 0.8. This results in disadvantageous increased noise, aerodynamic losses and possible mechanical excitation, phenomena which it is desirable to avoid or at least limit.
The present invention seeks to provide a nacelle profile that seeks to address the aforementioned problems.
Accordingly the present invention provides a propeller gas turbine engine comprising propellers and a nacelle, the nacelle comprising a forebody located upstream of the propellers and an afterbody located downstream of the forebody, the forebody comprising a first, upstream region and a second, downstream region, the first region having a convex profile including a maximum diameter intermediate its ends and the second region having a concave profile including a local minimum diameter. Advantageously, this shape presents a more diffused airflow to the propeller rotor stages, thereby reducing the aerodynamic losses and resultant noise compared to the prior art.
The nacelle may form a body of revolution or may be non-symmetrical about an axis of rotation of the gas turbine engine.
The nacelle may comprise G3 continuity or C3 continuity to present a smoother external surface to the airflow past the nacelle. This further reduces noise and aerodynamic losses.
The first and second regions meet at an intermediary location. The intermediary location may be positioned from one quarter to three quarters of the distance along the forebody in the downstream direction. Preferably the intermediary location is positioned from half to three quarters of the distance along the forebody in the downstream direction. This allows a smooth change in curvature between the first and second regions.
A maximum diameter of the nacelle may be located from one quarter to half of the distance along the forebody in the downstream direction. This also allows a smooth change in curvature between the first and second regions.
A minimum diameter of the nacelle may be coincident with an air intake at the upstream end of the forebody. Alternatively, the minimum diameter of the nacelle may be located downstream of the maximum diameter of the nacelle.
The diameter of the nacelle at each end of the forebody may be similar, a maximum diameter and a minimum diameter being located intermediate the ends of the nacelle.
The present invention also provides a propeller gas turbine engine, particularly a contra-rotating propeller gas turbine engine, comprising a nacelle as described.
The present invention will be more fully described by way of example with reference to the accompanying drawings, in which:

Figure 1 is a sectional side view of a conventional gas turbine engine having contra-rotating propeller stages.
Figure 2 is a schematic side view of a gas turbine engine having contra-rotating propeller stages according to the present invention.
Figure 3 is a graphical representation of the variation in |C_p| with distance along the forebody of a gas turbine engine according to the present invention.

An exemplary embodiment of the present invention is shown in Figure 2. A propeller gas turbine engine 10 is indicated having a rotational axis 9. The engine 10 has an air intake 12, compressor stages 14, 15, combustion equipment 16, turbine stages 17, 18, a power turbine 19 and a core exhaust nozzle 20. Front and rear propeller stages 23, 24 are shown towards the rear of the engine 10, which is in the pusher configuration. A nacelle 21 surrounds the core engine 11.
The nacelle 21 comprises a forebody 28 that extends between the air intake 12 and the rotor stage 23, and an afterbody 30 that extends between the forebody 28 and the core exhaust nozzle 20 of the engine 10, as indicated by labelled double-ended arrows on Figure 2. The nacelle 21 profile is preferably designed to have parametric C3 continuity, meaning that the profile when parameterised in terms of a parameter p has continuity of the rate of change of curvature at each connection point. This necessarily implies that the profile also has geometric G3 continuity of the rate of change of curvature at connection points. As is well understood in the field of parametric curve design, G3 continuity is not sufficient to imply C3 continuity but designing to each of G3 and C3 continuity necessarily includes the design achieving lower orders of continuity (G2, C2: continuity of curvature; G1, C1: continuity of tangency; G0, C0: continuity of connection).
The advantage of designing to G3 continuity or to C3 continuity, without presenting any local oscillations in the curvature, is that the pressure coefficient distribution across the nacelle 21 is smooth. This means that the velocity distribution does not present sudden variations and thereby avoids localised zones of high speed flow and shock waves, which are often present in conventional propeller gas turbine engines 10 flying at high speed. Further advantages accrue because aerodynamic losses are reduced and therefore fuel consumption is also reduced. Additionally, the smoother velocity and pressure coefficient distributions enable either higher flight speeds or lower aerodynamic losses or a combination of these, which are beneficial to airlines and popular with customers.
The forebody 28 of the nacelle 21 comprises two regions, a first upstream region 32 extending between the air intake 12 and a second downstream region 34, the second region extending between the first region 32 and the propeller stage 23. When considered as a section through the nacelle 21, the nacelle profile has the form of a line and the first and second regions 32, 34 of the nacelle 21 profile meet at an intermediary location 36. The three-dimensional nacelle 21 therefore comprises a ring or annulus of intermediary locations 36.
The nacelle 21 of the present invention solves the high speed problems of the prior art arrangements by arranging the forebody 28 to have a convex-concave profile. A convex portion includes a local maximum diameter whilst a concave portion includes a local minimum diameter. Preferably a convex portion has the local maximum diameter between its ends and a concave portion has the local minimum diameter between its ends. Thus the first region 32 has convex profile whilst the second region 34 has concave profile so that the airflow over the external surface of the nacelle 21 is diffused in the region immediately preceding the propeller stages 23, 24. Diffusion of the airflow in this manner reduces its velocity compared to the prior art arrangement thereby reducing the aerodynamic losses caused by the prior art arrangement and improving the fuel consumption of the engine 10 as a result. The intermediary location 36 between the first and second regions 32, 34 is defined as the location at which the curvature of the nacelle 21 profile is zero, the profile being parallel to the tangent. At this location the curvature changes sign from positive to negative or vice versa and the radius of curvature is infinite.
The intermediary location 36 is positioned between a quarter and three-quarters of the distance along the forebody 28 from the intake 12 towards the propeller stages 23, 24. This allows a smooth rate of change of curvature of the forebody 28 profile. In preferred embodiments, the intermediary location 36 is between half and three quarters of the distance along the forebody 28 and in a particularly preferred embodiment, the intermediary location 36 is located three quarters of the distance along the forebody 28 from the intake 12 towards the propeller stage 23. Where the maximum diameter of the nacelle 21 is located towards the intake 12, for example around a quarter of the distance along the forebody 28 from the intake 12, the intermediary location 36 is downstream from the maximum diameter. In this case the diameter of the forebody 28 at the intake 12 may be approximately the same diameter as at the end of the forebody 28 or may be larger or smaller depending on the specific application. Conversely, where the maximum diameter of the nacelle 21 is located close to the propeller stages 23, 24 the intermediary location 36 is upstream from the maximum diameter.
Figure 3 is a graphical representation of the magnitude of Cp against the distance along the forebody. Cp is a non-dimensional pressure coefficient describing the relative pressures of the air flow field. For the majority of the forebody 28, the pressure coefficient Cp is negative. The origin of the x-axis corresponds to the air intake 12 of the propeller gas turbine engine 10. Curve 38 corresponds to a conventional convex forebody 28 of the nacelle 21. The magnitude of the pressure coefficient Cp falls steeply as the air interacts with the air intake 12, which may be annular or a pitot intake, and then rises relatively steeply from the point at which the pressure coefficient Cp becomes negative. The pressure coefficient Cp then remains approximately constant for all positions along the forebody 28 towards the propeller stages 23, 24. If the maximum diameter of the nacelle 21 is larger than the diameter of the nacelle 21 at the hub of the propeller stages 23, 24, the magnitude of the pressure coefficient Cp may fall slightly towards the propeller stages 23, 24.
In contrast, the curve 40 corresponds to the nacelle 21 according to the present invention having the first region 32 being convex and the second region 34 being concave. The magnitude of the pressure coefficient Cp drops steeply and then rises relatively steeply from the point at which the pressure coefficient Cp becomes negative, as for curve 38 as the air interacts with the exterior of the air intake 12. The intake 12 may be annular or a pitot intake as for the prior art. However, the magnitude of the pressure coefficient Cp then decreases significantly as the forebody 28 changes from the convex first region 32 into the concave second region 34. Thus, consistent diffusion of the airflow past the external surface of the nacelle 21 occurs in advance of the first propeller stage 23, thereby reducing the speed of the airflow as it enters the rotors 23, 24, particularly the root portion thereof.
The diffusive nacelle 21 according to the present invention may comprise a body of revolution about the rotational axis 9 of the propeller gas turbine engine 10. Thus the nacelle 21 profile is rotated about the axis 9 to form a surface of revolution. In this case, the intermediary locations 36 of the profile between the first and second regions 32, 34 form an annulus in a plane that perpendicularly bisects the rotational axis 9. Alternatively, the nacelle 21 may be at least partially asymmetrical about the axis 9 provided that the first region 32 of the forebody 28 is convex and the second region 34 is concave. In this case, the intermediary locations 36 of the profile form a ring that may be deformed from circular in one or more dimensions.
Although the present invention has been envisaged for a nacelle 21 of a propeller gas turbine engine 10 mounted via a pylon from an aircraft wing or tail part, the principles are also applicable to engines 10 that are integrated within a wing or tail structure to further reduce aerodynamic losses.

Claims

A propeller gas turbine engine (10) comprising propellers (23, 24) and a nacelle (21), the nacelle (21) comprising a forebody (28) located upstream of the propellers (23, 24) and an afterbody (30) located downstream of the forebody (28), the forebody (28) comprising a first, upstream region (32) and a second, downstream region (34), the first region (32) having a convex profile including a maximum diameter intermediate its ends and the second region (34) having a concave profile including a local minimum diameter.
A propeller gas turbine engine (10) as claimed in claim 1 wherein the nacelle (21) forms a body of revolution about a longitudinal axis.
A propeller gas turbine engine (10) as claimed in claims 1 or 2 wherein the nacelle (21) comprises G3 continuity.
A propeller gas turbine engine (10) as claimed in any of claims 1 to 3 wherein the nacelle (21) comprises C3 continuity.
A propeller gas turbine engine (10) as claimed in any preceding claim wherein the first and second regions (32, 34) meet at an intermediary location (36).
A propeller gas turbine engine (10) as claimed in claim 5 wherein the intermediary location (36) is positioned from one quarter to three quarters of the distance along the forebody (28) in the downstream direction.
A propeller gas turbine engine (10) as claimed in claim 5 or 6 wherein the intermediary location (36) is positioned from half to three quarters of the distance along the forebody (28) in the downstream direction.
A propeller gas turbine engine (10) as claimed in any preceding claim wherein a maximum diameter of the nacelle (21) is located from one quarter to half of the distance along the forebody (28) in the downstream direction.
A propeller gas turbine engine (10) as claimed in any preceding claim wherein a minimum diameter of the nacelle (21) is coincident with an air intake (12) at the upstream end of the forebody (28).
A propeller gas turbine engine (10) as claimed in claim 8 wherein a minimum diameter of the nacelle (21) is located downstream of the maximum diameter of the nacelle (21).
A propeller gas turbine engine (10) as claimed in any preceding claim wherein the diameter of the nacelle (21) at each end of the forebody (28) is similar, a maximum diameter and a minimum diameter being located intermediate the ends of the nacelle (21).