Aircraft engine

Abstract

An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S.sub.tip, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition M.sub.rel, wherein M.sub.rel is between 0.4 and 0.93, L/D is between 0.2 and 0.45 and S.sub.tip is from −1 to 0.1.

Claims

1. An aircraft engine comprising: a fan, the fan having a diameter D, the fan including a plurality of fan blades, each fan blade having a tip sweep metric S.sub.tip and a leading edge, a forward-most portion on the leading edge of each fan blade being in a first reference plane; and a nacelle, comprising an intake portion upstream of the fan, a forward-most edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane; wherein the aircraft engine has a cruise design point condition M.sub.rel from 0.4 to 0.93, L/D is from 0.2 to 0.45 and S.sub.tip is from −1 to 0.1.

2. The aircraft engine of claim 1, wherein each fan blade has a tip sweep metric S.sub.tip from −0.75 to 0.1.

3. The aircraft engine of claim 1, wherein each fan blade has a tip sweep metric S.sub.tip between −0.5 and zero.

4. The aircraft engine of claim 1, wherein each fan blade has a tip sweep metric S.sub.tip greater than −0.25 and less than zero.

5. The aircraft engine of claim 1, wherein each fan blade has a tip sweep metric S.sub.tip greater than −0.1 and less than zero.

6. The aircraft engine of claim 1, wherein L/D is from 0.25 to 0.40.

7. The aircraft engine of claim 1, wherein L/D is from 0.30 to 0.40.

8. The aircraft engine of claim 1, wherein M.sub.rel from 0.4 to 0.85.

9. The aircraft engine of claim 1, wherein M.sub.rel from 0.4 to 0.80.

10. The aircraft engine of claim 1, wherein the aircraft engine is a gas turbine engine.

11. The aircraft engine of claim 10, wherein the gas turbine engine is a geared gas turbine engine.

12. The aircraft engine of claim 10, wherein the aircraft engine accommodates a first rate of air flow into an intermediate pressure compressor and a second rate of air flow which passes through a bypass duct, the ratio of the second rate of air flow to the first rate of air flow being greater than 10.

13. The aircraft engine of claim 1, wherein the aircraft engine is powered by electricity.

14. The aircraft engine of claim 1, wherein the fan diameter D is from 220 cm to 380 cm.

15. An aircraft having at least one aircraft engine according to claim 1.

Description

DESCRIPTION OF THE DRAWINGS

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

(2) FIG. 1 is a sectional side view of an example gas turbine aircraft engine;

(3) FIG. 2 is a second sectional side view of an example aircraft engine;

(4) FIG. 3 is a sectional sideview of a fan blade;

(5) FIG. 4 is a sectional plan view of a fan blade; and

(6) FIG. 5 is a second sectional plan view of a fan blade.

DETAILED DESCRIPTION

(7) Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art.

(8) With reference to FIG. 1, an example of an aircraft engine, in this case a gas turbine engine, is generally indicated at 10, having a principal and rotational axis 11. The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, an intermediate pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, an intermediate pressure turbine 18, a low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 generally surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.

(9) The aircraft engine 10 works in the conventional manner for a gas turbine engine so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the intermediate pressure compressor 14 and a second air flow which passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the air flow directed into it before delivering that air to the high pressure compressor 15 where further compression takes place. As such, the region of the engine with the intake portion 12 shall be referred to as the “front” or “forward” part of the engine, and the region of the engine with the exhaust nozzle 20 shall be referred to as the “back” or “rearward” part of the engine. This “front” and “back” nomenclature will be true for all types of engine considered here, including electric engines.

(10) Furthermore, with reference to all examples herein, the relative positions of components within the engine may be described in relation to the order in which air entering the engine flows over them. Air enters the engine at the intake, and those components or parts of components at which the air arrives first can be described as upstream or forward of those components that air arrives at later, which by comparison are downstream or backwards. For example, the front or forward-most part of a component is that part of the component that air arrives at first when travelling through an engine. Equally, the back or rear-most part of a component is that part of a component the air arrives at last when travelling through the engine. The front or forward-most part of a component is therefore upstream from the back or rear-most part of the component, and equally the back or rear-most part of the engine is downstream from the front or forward-most part of a component.

(11) 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 17, intermediate 18 and low-pressure 19 turbines before being exhausted through the exhaust nozzle 20 to provide additional propulsive thrust. The high 17, intermediate 18 and low-pressure 19 turbines drive respectively the high-pressure compressor 15, intermediate-pressure compressor 14 and fan 13, each by suitable interconnecting shaft. An alternative design of engine (not shown) is one that includes a gearbox between the fan 13 and one or more of the turbines 18, 19, usually positioned in the engine core just behind the fan, that allows the fan 13 to rotate at a different speed that the turbine driving it. This can be useful as it allows the fan and the turbine(s) to each rotate at their own optimised operating speeds during use. Such engines can be referred to as geared gas turbine engines.

(12) In the case of an electric engine, the fan 13 is powered by an electric motor, which can occupy a similar position behind the fan to the compressor section of the gas turbine engine. Such an engine can, for example, be powered by batteries stored elsewhere within the aircraft.

(13) FIG. 2 shows an exemplary layout of the fan and nacelle for an aircraft engine 10 with a selection of key parameters marked. The fan 13 comprises a number of fan blades 50, such as that shown in FIG. 3. The radius of the fan 13, also referred to as the fan tip radius, may be measured between the principle engine axis 11 and the tip r.sub.tip (see FIG. 3) of a fan blade 50 at its forward edge. The fan diameter (D) may simply be defined as twice the radius of the fan 13. The forward-most portion of the fan, i.e. the leading edges of the fan blades, are situated in a first reference plane 30. The nacelle 21, comprising the intake portion 12 forward of the fan 13, has forward edges 40 on the inlet portion which are situated in a second reference plane 32. The distance between the first reference plane 30 and second reference plane 32 is equal to the length L. This distance L is commonly referred to as the intake length.

(14) FIG. 3 shows a side view of a fan blade 50 with a tip sweep in the rotational axis, r. This view presents two parameters, x and r, needed to calculate the sweep of the fan blade tip, S.sub.tip. The x-direction is parallel to the principal and engine rotation axis 11 and perpendicular to r. A third parameter, relating to a leading edge line v, is shown in FIG. 5. The extent along the blade in the radial direction is denoted by r, extending from r=0.0 or r.sub.hub (i.e. the base of the blade at its leading edge) to r=1.0 or r.sub.tip (i.e. the furthest radial extent of the blade at its leading edge), the distance from 0 to 1 being equal to the span of the blade. As an example, r.sub.0.7 is located at 70% along the span from the hub to the tip. x represents the position of the leading (forward) edge of the blade in the plane perpendicular to r and parallel to the principal and rotational axis 11, with x being positive as you move from the front to the back of the engine, as indicated by the axis arrow in FIG. 3. It is important to note that the blade does not need to have a shape which is negative in the x-axis in order to have a negative tip sweep metric. S.sub.tip; a neutral or positive curve in the x-axis can be complimented with a negative curve in the leading edge line v-axis to produce a blade tip with a small positive or overall negative tip sweep metric.

(15) The section shown in FIG. 4 represents a cross-sectional plane through the fan blade 50 at an arbitrary span fraction location. The cross section of the blade has a camber line 52. In general, the camber line 52 may be defined as a line that is equidistant from a pressure surface 56 and a suction surface 54 for that cross-section. The cross section is characterised with a leading edge 60, and trailing edge 58. The leading edge angle, χ, may be defined as the angle the camber line 52 makes with respect to the engine rotation axis x at the leading edge 60 of the cross-section. In other words, the leading edge angle χ is the angle between the engine rotation axis x and the line normal to the leading edge 60 at the point where the camber line 52 meets the leading edge 60. A positive angle may be defined measuring anti-clockwise from the engine axis. A negative angle may be defined measuring clockwise from the engine axis. It is possible χ may vary from positive to negative along the blade span. In the example illustrated in FIG. 4, χ is positive.

(16) FIG. 5 shows two cross-sections at r.sub.0.9 (grey) 100 and r.sub.1.0 (black) 200 span fraction. The cross-sections have respective suction sides 54a,54b, pressure sides 56a,56b, trailing edges 58a,58b and leading edges 60a,60b. The r.sub.1.0 section 200 has been displaced relative to the r.sub.0.9 section 100. Along the engine rotation axis x, the r.sub.1.0 section 200 has been displaced by x.sub.1.0−x.sub.0.9. Following the convention defined, x.sub.1.0−x.sub.0.9 illustrated in FIG. 5 is a negative value. In direction of rotation the r.sub.1.0 section 200 has been displaced by r.sub.1.0θ.sub.1.0−r.sub.0.9θ.sub.0.9. Following the convention defined, r.sub.1.0θ.sub.1.0−r.sub.0.9θ.sub.0.9 illustrated in FIG. 5 is a negative value.

(17) A leading edge line 53 is shown, which is the line normal to the leading edge 60a at the point where the camber line (not shown) meets the leading edge 60a of the r.sub.0.9 section 100. The angle the leading edge line 53 makes with the engine rotation axis x at the leading edge 60a of the r.sub.0.9 section 100 is equal to χ.sub.0.9, as shown. The r.sub.1.0 section 200 is displaced parallel to the leading edge line by v.sub.1.0−v.sub.0.9. The value of v.sub.1.0−v.sub.0.9 can be calculated using the following formula:
v.sub.1.0−v.sub.0.9=√{square root over ((x.sub.1.0−x.sub.0.9).sup.2+(r.sub.1.0θ.sub.1.0−r.sub.0.9θ.sub.0.9).sup.2)} cos(χ.sub.0.9−a tan 2(r.sub.1.0θ.sub.1.0−r.sub.0.9θ.sub.0.9,x.sub.1.0−x.sub.0.9))

(18) Where atan2(r.sub.1.0θ.sub.1.0−r.sub.0.9θ.sub.0.9,x.sub.1.0−x.sub.0.9) is the four-quadrant inverse tangent defined:

(19) $\mathrm{atan}2\left(r_{1.0}{\theta }_{1.0}-r_{0.0}{\theta }_{0.9},x_{1.0}-x_{0.9}\right)=\{\begin{array}{cc}{\tan }^{-1}\left(\frac{r_{1.0}{\theta }_{1.0}-r_{0.9}{\theta }_{0.9}}{x_{1.0}-x_{0.9}}\right),& \left(x_{1.0}-x_{0.9}\right)>0\\ {\tan }^{-1}\left(\frac{r_{1.0}{\theta }_{1.0}-r_{0.9}{\theta }_{0.9}}{x_{1.0}-x_{0.9}}\right)+\pi ,& \left(x_{1.0}-x_{0.9}\right)<0\mathrm{and}\left(r_{1.0}{\theta }_{1.0}-r_{0.9}{\theta }_{0.9}\right)\ge 0\\ {\tan }^{-1}\left(\frac{r_{1.0}{\theta }_{1.0}-r_{0.9}{\theta }_{0.9}}{x_{1.0}-x_{0.9}}\right)-\pi ,& \left(x_{1.0}-x_{0.9}\right)<0\mathrm{and}\left(r_{1.0}{\theta }_{1.0}-r_{0.9}{\theta }_{0.9}\right)<0\\ +\frac{\pi }{2},& \left(x_{1.0}-x_{0.9}\right)=0\mathrm{and}\left(r_{1.0}{\theta }_{1.0}-r_{0.9}{\theta }_{0.9}\right)>0\\ -\frac{\pi }{2},& \left(x_{1.0}-x_{0.9}\right)=0\mathrm{and}\left(r_{1.0}{\theta }_{1.0}-r_{0.9}{\theta }_{0.9}\right)<0\\ 0& \left(x_{1.0}-x_{0.9}\right)=0\mathrm{and}\left(r_{1.0}{\theta }_{1.0}-r_{0.9}{\theta }_{0.9}\right)=0\end{array}$

(20) Following the convention defined, v.sub.1.0−v.sub.0.9 illustrated in FIG. 5 is a negative value.

(21) Using these terms, the sweep of the fan blade 50 can be defined in terms of

(22) $\frac{\mathrm{dv}}{\mathrm{dr}},$
referred to as the span stack slope.

(23) In this region, i.e. in the top 10% of the radial blade span, we define a tip sweep metric, S.sub.tip, as the span stack slope

(24) $\frac{\mathrm{dv}}{\mathrm{dr}}$
between 0.9 (90%) and 1.0 (100%) of the span, normalised by a proxy of the relative Mach number at cruise design point conditions (i.e. the relative magnitude of the combined vectors of the velocity of the blade and the velocity of the air it is passing through compared with the speed of sound under the same conditions), M.sub.rel:

(25) $S_{\mathrm{tip}}=\frac{{\left(\frac{\mathrm{d\nu }}{\mathrm{dr}}\right)}_{0.9-1.0}}{M_{\mathrm{rel}}}$

(26) Where

(27) ${\left(\frac{\mathrm{dv}}{\mathrm{dr}}\right)}_{0.9-1.0}$
is defined as:

(28) ${\left(\frac{\mathrm{dv}}{\mathrm{dr}}\right)}_{0.9-0.10}=\frac{\left(\frac{v_{1.0}-v_{0.9}}{r_{1.0}-r_{0.0}}\right)}{1.0-0.9}$

(29) Where M.sub.rel is defined at the cruise design point condition (modulus function applied so the value of M.sub.rel is always positive):

(30) $M_{\mathrm{rel}}=\sqrt{{\left(\frac{\left(\frac{\tilde{r}\Omega }{\sin \tilde{\chi }}\right)}{\sqrt{1.4\times 287\times 250.5}}\right)}^2}$

(31) Where Ω=cruise design point rotational speed in radians per second, and where {tilde over (χ)} is the discrete area-averaged leading edge angle:

(32) $\tilde{\chi }=\frac{\begin{array}{c}{\chi }_{0.1}\pi \left(r_{0.2}^2-r_{0.0}^2\right)+{\chi }_{0.3}\pi \left(r_{0.4}^2-r_{0.2}^2\right)+\\ {\chi }_{0.5}\pi \left(r_{0.6}^2-r_{0.4}^2\right)+{\chi }_{0.7}\pi \left(r_{0.8}^2-r_{0.6}^2\right)+{\chi }_{0.9}\pi \left(r_{1.0}^2-r_{0.8}^2\right)\end{array}}{\pi \left(r_{1.0}^2-r_{0.0}^2\right)}$

(33) Where {tilde over (r)} is the discrete area-averaged rotor inlet radii:

(34) $\tilde{r}=\frac{\begin{array}{c}{r}_{0.1}\pi \left(r_{0.2}^2-r_{0.0}^2\right)+r_{0.3}\pi \left(r_{0.4}^2-r_{0.2}^2\right)+\\ r_{0.5}\pi \left(r_{0.6}^2-r_{0.4}^2\right)+r_{0.7}\pi \left(r_{0.8}^2-r_{0.6}^2\right)+r_{0.9}\pi \left(r_{\mathrm{tip}}^2-r_{0.8}^2\right)\end{array}}{\pi \left(r_{1.0}^2-r_{0.0}^2\right)}$

(35) In an engine which has a small L/D ratio and operates at low M.sub.rel employing fans with a tip sweep metric S.sub.tip from −1 to 0.1 has an advantageous effect on the ability to resist rotating stall of the main fan after the air flow entering the nacelle has separated from the intake due to a high angle of attack. The degree of the effect is such that it justifies the increased manufacturing challenges and costs involved.

(36) There is a benefit to shaping the sweep of the fan blade for an engine which has a small L/D ratio, and operates at low M.sub.rel by having a tip sweep metric with S.sub.tip having a value from −1 to 0.1. Therefore, contrary to established belief, a low-speed fan can benefit from being swept, particularly with a tip sweep metric S.sub.tip from −1 to 0.1, if the low-speed fan is part of an aircraft engine with an intake L/D ratio of between 0.2 and 0.45 operating at low M.sub.rel.

(37) Other aircraft engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan. Alternatively, the aircraft engine may be powered by electricity, i.e. it may be an electric aircraft engine.

(38) It will be understood that the disclosure 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 sub-combinations of one or more features described herein.