Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle defining a duct, and a fan located within the duct. The fan comprises a hub arranged to rotate around a rotational axis (A-A) and a plurality of blades attached to the hub, each of which has a span from a root at the hub defining a 0 percent span position (rhub) to a tip defining a 100 percent span position (rtip) and a plurality of span positions therebetween (r?[rhub, rtip]). A hub-tip ratio of the fan, defined as the ratio of the diameter of the hub to the diameter of the fan measured at the leading edge of the blades, is from 0.45 to 0.55.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle (501) defining a duct (502), and a fan (503) located within the duct. The fan comprises a hub arranged to rotate around a rotational axis (A-A) and a plurality of blades attached to the hub. Each blade has a span (r) from a root at the hub defining a 0 percent span position (r=0) to a tip defining a 100 percent span position (r=1) and a plurality of span positions therebetween (r ? [0, 1]), and a stagger angle at the 0 percent span position (?hub) relative to the rotational axis of 40 degrees or greater.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft includes a nacelle (501) defining a duct (502), and a fan (503) located therewithin. The fan comprises a hub which rotates around a rotational axis (A-A) and a plurality of blades attached thereto. Each blade has a span (r) from a root at the hub defining a 0 percent span position (r=0) to a tip defining a 100 percent span position (r=1) and a plurality of span positions therebetween (r ?[0, 1]), a leading edge and a trailing edge defining, for each span position, a chord therebetween having a chord length (c), and a blade thickness (t) defined for each span position thereof. For each blade, a ratio of thickness at the 0 percent span position (thub) to chord length is 0.1 or greater.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle (501) defining a duct, and a fan located therewithin. The fan comprises a hub arranged to rotate around a rotational axis (A-A) and a plurality of blades attached thereto. Each blade has a span (r) from a root at the hub defining a 0 percent span position (r=0) to a tip defining a 100 percent span position (r=1) and a plurality of span positions therebetween (r ? [0, 1]), and leading and trailing edges defining, for each span position, a chord therebetween to having a chord length (c). For each of said plurality of blades, the ratio of chord length at the 0 percent span position (chub) to chord length at the 100 percent span position (ctip) is 1 or greater.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle defining a duct, and a fan located within the duct. The fan comprises a hub arranged to rotate around a rotational axis and a plurality of blades attached to the hub, each of which has a span from a root at the hub defining a 0 percent span position (rhub) to a tip defining a 100 percent span position (rtip) and a plurality of span positions therebetween (r?[rhub, rtip]). The hub has a negative hade angle (?) with respect to the rotational axis at an axial o position coincident with the leading edge of the blades.

Abstract: A boundary layer ingestion fan system for location aft of the fuselage of an aircraft is shown. It comprises a nacelle defining a duct, and a fan located within the duct. The fan comprises a hub arranged to rotate around a rotational axis (A-A) and a plurality of blades attached to the hub. A blade blockage, which is the ratio of the blade thickness to the product of the circumferential pitch and the cosine of a blade inlet angle (t/s·cos?1), is 0.25 or greater at the 0 percent span position.

Abstract: A turbomachine (105) configured to compress supercritical carbon dioxide is shown. The turbomachine comprises, in fluid flow series, an inlet (201), an inducerless radial impeller (202) having a plurality of blades, and a fully vaneless diffuser (203). The radius of the inlet (r0) is from 25 to 50 percent of the radius of the impeller (r2).

Abstract: A turbomachine (105) configured to compress supercritical carbon dioxide is shown. The turbomachine comprises, in fluid flow series, an inlet (201), an inducerless radial impeller (202) having a plurality of blades, a fully vaneless diffuser (203), and a volute (204) comprising a tongue and having a flow area at the tongue equal to that of the diffuser.

Abstract: A turbomachine (105) configured to compress supercritical carbon dioxide is shown. The turbomachine comprises, in fluid flow series, an inlet (201), an inducerless radial impeller (202) having a plurality of blades, and a fully vaneless diffuser (203).

Abstract: A turbomachine (105) configured to compress supercritical carbon dioxide is shown. The turbomachine comprises, in fluid flow series, an inlet (201), an inducerless radial impeller (202) having a plurality of blades, and a fully vaneless diffuser (203). The inlet is radially flared to induce a radial component in flow prior to an entry to the impeller.

Abstract: A turbomachine (105) configured to compress supercritical carbon dioxide is shown. The turbomachine comprises, in fluid flow series, an inlet (201), an inducerless radial impeller (202) having a plurality of backswept blades (211,212) each of which have a blade exit angle (?2) of from ?50 to ?70 degrees, and a fully vaneless diffuser (203).

Abstract: A boundary layer propulsor comprises a rotor and a plurality of first aerofoil blades. The rotor has an axis of rotation. The plurality of first aerofoil blades extends radially from the rotor and is arranged in a circumferential array around the axis of rotation. Each of the first aerofoil blades has, in a radially outward sequence, a radially proximal portion, a middle portion, and a radially distal portion. The radially proximal portion has a first cambered cross-section, the middle portion has a second uncambered cross-section, and the radially distal portion has a third cambered cross-section. The first cambered cross-section is cambered in an opposite sense to the third cambered cross-section.

Abstract: A method for defining the shape of an edge of an aerofoil comprises the following steps: obtaining a starting value for the thickness of the edge; defining a distance from the edge of the aerofoil in terms of the thickness; over the defined distance, defining a space S in terms of normalised camberline and normalised thickness; transforming the space S to a new space S?; parameterising the new space S? to obtain new values for normalised camberline and normalised thickness; defining a new shape for the edge using the new parameter values. In a preferred embodiment, the transforming step comprises applying a parabolic function to the space S.