103results about "Heat reducing structures" patented technology

An aircraft includes a wing and a leading-edge aerodynamic surface coupled to the wing. An actuator is disposed within the wing and coupled to the leading-edge aerodynamic surface via a linkage system. Two or more hinge systems are coupled to the leading-edge aerodynamic surface and configured to rotate about a fixed axis internal to the wing to move the leading-edge aerodynamic surface between a retracted position and a drooped position.

The invention discloses a low-frequency broadband sound absorption metamaterial adaptive wing and a design method thereof, and belongs to the technical field of aircraft design. The problem that in the prior art, a traditional aircraft wing and a design method thereof are difficult to achieve the low-frequency and broadband noise reduction effect is solved. The wing comprises a leading edge inner side slat, a leading edge outer side slat, a main wing, a trailing edge inner side flap and a trailing edge outer side flap, the leading edge inner side slat and the leading edge outer side slat are connected with the front side of the main wing, and the trailing edge inner side flap and the trailing edge outer side flap are connected with the rear side of the main wing; the metamaterial sound absorption strips are embedded into concave cavities of the leading edge inner side slat and the leading edge outer side slat respectively, the metamaterial sound absorption strips are embedded into the leading edge of the main wing, each metamaterial sound absorption strip is composed of a Helmholtz resonator array composed of single-layer Helmholtz resonator units, and each metamaterial sound absorption strip is provided with a micro-perforated surface with continuous curvature. The noise reduction level of the wing is effectively improved, the sound absorption frequency band is wider, and the method can be applied to sound absorption wing design.

An aircraft flap comprises an upper wing surface and a lower wing surface which are opposite to each other, and a cavity is defined by the upper wing surface and the lower wing surface; the noise reduction device is arranged in the cavity and comprises a first noise reduction structure, the first noise reduction structure comprises a perforated plate and an upper part and a lower part, the upper part is embedded in the upper wing surface, and the lower part is embedded in the lower wing surface; a second noise reduction structure including foam metal, disposed between the upper portion and the lower portion of the first noise reduction structure, and attached to the first noise reduction structure; and the third noise reduction structure comprises a super-structure partition plate and is embedded into the second noise reduction structure. According to the composite superstructure design integrating the three noise reduction structures together, the structural stability of the noise reduction device can be guaranteed, and meanwhile, the jet flow installation noise of the aircraft engine can be effectively reduced.

The invention relates to the technical field of aerospace, and particularly provides a liquid metal closed-loop thermal management system and method for a hypersonic aircraft. The system comprises a liquid metal circulation loop which comprises a liquid metal pump used for driving a liquid metal working medium to circularly flow; the heat absorption device is integrated in the liquid metal circulation loop and used for receiving operation heat of the hypersonic flight vehicle and transmitting the operation heat to the liquid metal working medium flowing through the hypersonic flight vehicle. The thermo-acoustic power generation device at least comprises a hot end heat exchanger connected to the heat absorption downstream of the liquid metal circulation loop in series and used for converting part of heat energy of the liquid metal working medium heated after heat absorption into electric energy so as to drive a liquid metal pump. The liquid metal circulation loop further comprises a first diverter valve and a second diverter valve, the first diverter valve is connected into the heat absorption device, and the second diverter valve is connected into the thermo-acoustic power generation device. The dynamic cooperative controller is used for controlling the dynamic cooperative controller based on the flight state parameters of the hypersonic aircraft, the engine thrust requirement and the temperature of the key thermal protection part. The total flow of the liquid metal working medium and flow distribution in the heat absorption device and the thermo-acoustic power generation device are dynamically and cooperatively adjusted by controlling the liquid metal pump, the first diverter valve and the second diverter valve. According to the invention, the compactness of the system can be obviously improved, the heat sink consumption is greatly reduced, and a reliable and efficient heat management solution is provided for long-endurance hypersonic flight.

The invention discloses a variable-pore phase-change sweating cooling bionic structure based on fractal characteristics. The variable-pore phase-change sweating cooling bionic structure comprises a variable-pore porous sweating layer and a regulation and control unit, wherein the variable-pore porous sweating layer is composed of fractal flow equalizing plates which are sequentially arranged in a stacked mode in the heat flow direction; the fractal flow equalizing plate is of an N-stage bifurcated tree-shaped structure and is used for equally dividing one path of cooling working medium into 2N paths of cooling working medium, and N is an integer ranging from 4 to 6; the variable-pore porous sweating layer is divided into an upper-layer high-heat-flow area and a lower-layer low-heat-flow area in the thickness direction, the upper-layer high-heat-flow area is used for promoting phase change and efficient heat exchange of a cooling working medium, and the lower-layer low-heat-flow area is used for promoting liquid-phase permeation of the cooling working medium; and the regulation and control unit predicts the phase change leading edge position S (t) in real time based on a neural network, and dynamically regulates the cooling working medium flow Q (t) input into the inlet of the fractal flow equalizing plate according to the S (t). The invention further discloses a design method of the structure. By introducing the bionic variable porosity characteristic, coupling of flowing heat exchange is achieved, the cooling efficiency of thermal protection is improved, and consumption of a heat exchange working medium is reduced.

The invention relates to a hypersonic aircraft leading edge active and passive composite thermal protection structure and a preparation method thereof. The thermal protection structure comprises a high-temperature metal heat pipe, a parting flow channel layer plate and a porous sweating layer plate which are arranged in a stacked mode from inside to outside, the parting flow channel layer plate is provided with a branch-shaped parting network flow channel, and the porous sweating layer plate is provided with concave grooves which are distributed in a staggered mode in the shape of the Chinese character'yao '. And the bottom of the sweating groove is communicated with the dendritic parting network flow channel through the sweating holes. Passive heat dissipation of the high-temperature metal heat pipe and active cooling of the parting flow channel-porous sweating layer are used for cooperative work, and advantage complementation of passive heat protection and active heat protection is achieved; the parting runner-porous sweating layer design can realize uniform distribution of a coolant on the porous sweating layer plate, and a large number of micro-pore channels formed by the porous sweating layer micro pin fins-interconnected embedded grooves significantly promote sweating cooling, so that efficient active and passive composite thermal protection is realized.

A winglet system is configured to be secured to a wing of an aircraft. The winglet system includes an upper blade, a lower blade, and a cover plate integrally formed with the upper blade and the lower blade. The cover plate covers an opening within a transition joint between the upper blade and the lower blade. The cover plate remains coupled to the upper blade if the lower blade is separated from the winglet system.

The invention relates to a heated leading edge structure for an aircraft, a slat, a wing and an aircraft. A heated leading edge structure (14) for an aircraft (10) includes a leading edge panel (20) having an outer surface (32) configured for contact with an ambient flow (A) and an inner surface (34) opposite the outer surface (32); a reinforcing member (22) connected to the inner surface (34) of the leading edge panel (20); and a heating device (24) configured to heat the leading edge panel (20); the heating device (24) comprises at least one indirect heating device (44) mounted on the reinforcing member (22), said indirect heating device (44) being configured for heating the leading edge panel (20) at least by means of conduction through the reinforcing member (22).

The application discloses a kind of torsion wing structure and aircraft, it is related to the field of morphing wing technology, including linear drive device, wing rib and pressure-torsion coupling cell, wing rib and pressure-torsion coupling cell are multiple, multiple wing ribs are arranged along the wing span direction interval, and pressure-torsion coupling cell is equipped between two adjacent wing ribs;Pressure-torsion coupling cell includes intermediate connecting component and the load-bearing end block fixedly connected with the two ends of intermediate connecting component, load-bearing end block is fixedly connected with adjacent wing rib, and all wing ribs and pressure-torsion coupling cell form wing main body;The first end wing rib of wing main body is used to be fixedly connected with fuselage;Two load-bearing end blocks of pressure-torsion coupling cell are connected with linear drive device;Linear drive device can make the two load-bearing end blocks of each pressure-torsion coupling cell approach each other or away from each other, to make intermediate connecting component torsion deformation and drive wing rib around wing span torsion by load-bearing end block.The application has the effect of "lightweight-structure simplification-aerodynamic efficiency-high life-low maintenance cost".

Disclosed is a wing (3) for an aircraft (1), comprising a main wing (5) and a trailing edge high lift assembly (9) movably arranged at a trailing edge of the main wing (5), the trailing edge high lift assembly (9) comprising a flap (11) and a connection assembly movably mounting the flap (11) to the main wing (5), such that the flap (11) is movable between at least one retracted position (12, 14) and at least one extended position (15), wherein in the at least one extended position (15) a gap (16) is present between a lower rear panel (17) of the main wing (5) and a leading edge section (19) of the flap (11), and wherein in the at least one retracted position (12, 14) the gap (16) is closed or minimized. The object of the invention, to provide a wing that might be used as a morphing wing capable of adjusting wing area, curvature, and angle of attack during the entire flight, while having optimized aerodynamic performance throughout all flap settings, is achieved in that the flap (11) is movable between at least one extended position (15), a first retracted position (12), and a second retracted position (14) where the flap (11) is moved beyond the first retracted position (12) in a forward direction (13), and the wing comprises a gap closing device (23) configured to close or minimize the gap (16) in both the first retracted position (12) and the second retracted position (14) of the flap (11), and to open the gap (16) in the at least one extended position (15) of the flap.

The invention is notably directed to an aircraft, which includes an aircraft structure, a thermal exchange system, a thermal insulation material, and fasteners. The thermal exchange system includes a fluid-based cooling circuit and heat exchanger tiles. The latter cover, at least partly, the aircraft structure. They may thus form the skin of the aircraft. The fasteners secure the tiles to the aircraft structure. Each tile of the heat exchanger tiles embeds one or more fluidic channels, which are connected to the fluid-based cooling circuit, whereby the thermal exchange system is adapted to actively cool the tiles, in operation. Remarkably, the thermal insulation material extends between the aircraft structure and the heat exchanger tiles. The thermal insulation material is clamped on either side by the heat exchanger tiles and by the aircraft structure, thanks to the fasteners.

The method includes obtaining (402) thermal energy from a structure (112) to be cooled, the structure including microchannels (208). The method also includes supplying (402) the thermal energy to a water-based polymer network (206), the water-based polymer network including a gel formed using a polymer and water. The method further includes generating (404) one or more gases (210) by heating the water-based polymer network, the generating one or more gases including releasing water in the water-based polymer network to produce steam. Additionally, the method includes passing (408) the one or more gases through the microchannels to remove at least a portion of the thermal energy from the structure.

The invention discloses a distributed thermoelectric element variant wing SMA driver and method and application, the SMA driver comprises an active cooling layer, an SMA wire driving structure and an elastic supporting layer which are sequentially arranged from top to bottom, each layer is embedded in a flexible material in a sealed mode, the active cooling layer is composed of a plurality of thermoelectric elements arranged in an array mode, and the elastic supporting layer is embedded in the active cooling layer. The thermoelectric element achieves refrigeration of the SMA wire driving structure through the Peltier effect. The cooling efficiency can be remarkably improved, heat accumulation is eliminated, the recovery time of the driver is shortened, the driving frequency is improved, and the service life of the shape memory alloy driver is prolonged. According to the algorithm, an adaptive thermoelectric element layout scheme can be automatically generated in combination with the structural characteristics of the driver, so that heat transfer is accelerated, the uniformity of a temperature field is improved, the application requirements under multiple working conditions and complex environments are met, the algorithm has good universality, the integration requirements of different types of wings are met, and the application prospect is wide. The design process is effectively simplified, and the manufacturing cost is reduced.

The present invention relates to an aircraft (1002) comprising a wing (1001) having a fixed wing (1005) and a wingtip device (1003) movably mounted at the outer end of the fixed wing. The wingtip device (1003) is capable of moving between a flight configuration and a ground configuration about a rotation axis. The wingtip device (1003) abuts the fixed wing (1005) along an interface cut line (1035). The interface cut line (1035) includes a curved portion (1039) centered on the rotation axis (1011), the radius of the curved portion increasing continuously as the cut line moves about the axis (e.g., a spiral shape). The present invention also relates to a method of designing a wing, a fixed wing, a wingtip device, and an interface.

The invention is notably directed to an aircraft, which includes an aircraft structure, a thermal exchange system, a thermal insulation material, and fasteners. The thermal exchange system includes a fluid-based cooling circuit and heat exchanger tiles. The latter cover, at least partly, the aircraft structure. They may thus form the skin of the aircraft. The fasteners secure the tiles to the aircraft structure. Each tile of the heat exchanger tiles embeds one or more fluidic channels, which are connected to the fluid-based cooling circuit, whereby the thermal exchange system is adapted to actively cool the tiles, in operation. Remarkably, the thermal insulation material extends between the aircraft structure and the heat exchanger tiles. The thermal insulation material is clamped on either side by the heat exchanger tiles and by the aircraft structure, thanks to the fasteners.

This application discloses a foldable, tilted wingtip with an aerodynamic device. The disclosed exemplary wing for use with an aircraft includes a fixed portion and a folding portion proximate a distal end of the wing. The folding portion includes an inclined surface. The wing also includes at least one of a feather or a winglet, and a hinge operably coupled between the fixed portion and the folding portion to enable the folding portion to fold relative to the fixed portion.

This invention proposes an aerodynamic drag device and layout for a tailless aircraft. The aerodynamic drag device comprises three drag plates arranged side-by-side on the aircraft wing: an inner plate, a middle plate, and an outer plate. One side of each of the inner, middle, and outer plates is rotatably connected to the wing. The middle plate is located between the inner and outer plates, with the inner plate closer to the aircraft's centerline. The layout includes a blended fuselage and wing design, with a pair of ailerons and the aforementioned aerodynamic drag device symmetrically arranged at the wing's trailing edge. The opening angle of the pair of aerodynamic drag devices is independently controlled. This invention increases the aircraft's directional stability, reduces supersonic drag, improves supersonic handling performance, enhances cruise and maneuverability, and does not introduce additional adverse effects on the airframe structure.

The present invention provides a flap vortex disturbing device and a wing assembly (10) comprising the flap vortex disturbing device, which can disturb vortex generated at the inner and outer side end edges of a flap, thereby suppressing the reduction of lift force and the advance of stall of an aircraft, and a wing assembly (10) comprising the flap vortex disturbing device. The flap vortex interference devices (120 and 140) interfere vortexes of trailing edge flaps (210 and 220) arranged on the trailing edge of a fixed wing (100) when an aircraft flies at a high angle of attack, and the flap vortex interference devices are correspondingly arranged at the inner and outer side end edges (211 and 221) of the trailing edge flaps respectively and are in an obtuse triangle shape. And gaps for air flow to circulate are respectively formed between the flap vortex interference devices (120 and 140) and the inner and outer side end edges (211 and 221) of the trailing edge flap.

To provide fairings to support engines relative to wings of an aircraft, and methods for reducing drag.SOLUTION: A wing comprises a fairing 130 covering a pylon to connect to an engine. The fairing defines a horizontal plane, a first plane perpendicular to a longitudinal axis, and a second plane perpendicular to the longitudinal axis and aft of the first plane. A reference line is defined as the longitudinal axis when projected onto the horizontal plane. The fairing includes a fairing body defining an aerodynamic surface. The aerodynamic surface has an outboard portion and an inboard portion. The inboard portion is configured such that the first plane intersects the horizontal plane and the aerodynamic surface of the inboard portion at a first intersection point. The first intersection point is laterally displaced from the reference line by a first distance. The second plane intersects the horizontal plane and the aerodynamic surface of the inboard portion at a second intersection point. The second intersection point is laterally displaced from the reference line by a second distance, which is greater than the first distance.SELECTED DRAWING: Figure 2

The present application relates to a large-sized heat-resistant load-bearing integrated aircraft wing, belonging to the technical field of aircraft structures. The large-sized heat-resistant load-bearing integrated aircraft wing includes a fuselage shell. On both sides of the fuselage shell, a first installation groove and a second installation groove are provided. One end of the front wing facing the middle wing is provided with a first groove, and on the middle wing, there is a first lug that is inserted in cooperation with the first groove. One end of the middle wing facing the rear wing is provided with a second groove, and on the rear wing, there is a second lug that is inserted in cooperation with the second groove. The front wing, the middle wing, and the rear wing are connected between the first installation groove and the second installation groove through a front support and a rear support. Heat-resistant putty is arranged at the butt joint between the front wing, the middle wing, and the rear wing; in the present application, through the movable installation of the rear support, the connection structure between it and the fuselage shell can be thermally deformed and matched, avoiding the completely fixed installation method. When the wing generates thermal deformation, it cannot be extended synchronously, resulting in the wing being easily damaged.

The application discloses a hydrogen-electric hybrid energy supersonic passenger plane, and belongs to the field of engineering technology.The plane comprises an aerodynamic shape, a hydrogen-electric hybrid driven power device and an integral hydrogen tank for storing hydrogen fuel, wherein the aerodynamic shape comprises an elongated body, a wing containing a lifting surface, a tail wing and a canard; the power device comprises a gas turbine main engine and an electric propulsion ducted fan; the integral hydrogen tank comprises a pressure-bearing inner tank, an adiabatic layer and a support structure; the power device is connected with the body and the wing; and the hydrogen tank is connected with the body through the support structure.The scheme increases the lifting surface on the wing to ensure the high lift-drag ratio and low sound blast requirement of the plane, introduces hydrogen energy to realize zero carbon emission, and adopts the integral hydrogen tank to increase the hydrogen fuel storage of the plane and guarantee the voyage requirement of the plane.

The invention discloses a method for improving heat sink of an aircraft, which comprises the following steps of: arranging a cooling flow channel and an oil inlet pipe on an upper skin of a wing, inputting hot return oil to be cooled into the cooling flow channel through the oil inlet pipe, cooling the hot return oil by utilizing the heat sink of the upper skin, and arranging an oil conveying pipe at an outlet of the cooling flow channel and at a position close to a lower skin in a wing oil tank, the cooled hot return oil is mixed with original low-temperature fuel oil in the wing oil tank, and the mixed oil is conveyed to the oil collecting tank together and enters each heat dissipation system as a heat sink again to form circulation; high-temperature fuel oil is in large-area full contact with the wing upper skin through a cooling flow channel for heat exchange, high-temperature hot return oil is cooled through a low-temperature air heat sink outside the upper skin so as to improve the heat sink of the high-temperature hot return oil, the cooled return oil and low-temperature fuel oil in a wing oil tank are mixed and enter an oil collecting tank, the fuel oil is cooled again through an air heat sink of the wing lower skin, and the fuel oil is cooled again. The aircraft does not need to introduce ram air to cool the hot return oil, so that the fuel oil cooling efficiency and the aircraft stealth performance are improved, and the flight resistance is reduced.

The invention provides a wingtip vortex active control method and system based on surface pressure feedback and synthetic jet and a medium, and the method comprises the steps: collecting pressure data in real time based on a pressure sensor array arranged on the surface of a wing; performing time-frequency analysis on the pressure data, and extracting the shedding frequency and pressure pulsation characteristics of the current wingtip vortex; retrieving flow field characteristics based on the wingtip vortex optimal disturbance modal database of the pre-solution analysis to obtain dimensionless information; converting the dimensionless information into driving frequency, driving voltage and a phase instruction of a synthetic jet actuator to generate periodic jet; continuously monitoring the power spectrum density and suction peak change of a pressure sensor at a specific position of a wing, and judging whether a flow field meets a rapid instability condition or not; and if the rapid instability condition is not met, the frequency or phase difference of the synthetic jet is adjusted, rapid capture and closed-loop control from flow field data to optimal disturbance application can be achieved, and instability and rapid attenuation of the wingtip vortex are effectively promoted.

A heated leading edge structure (14) for an aircraft (10) including a leading edge panel (20) having an outer surface (32) configured to contact an ambient flow (A), and an inner surface (34) opposite the outer surface (32); a reinforcing member (22) connected to the inner surface (34) of the leading edge panel (20), and a heating arrangement (24) configured to heat the leading edge panel (20); wherein the heating arrangement (24) includes an indirect heating device (44) mounted on the reinforcing member (22) and configured to heat the leading edge panel (20) by conduction through the reinforcing member (22).

The invention discloses an ultrathin composite titanium-based material high-temperature-resistant and high-strength connecting structure and application. The connecting structure comprises a main bearing core body, a conical cushion block, an indexing thread locking block, an indexing inclined cushion block, a locking nut and an elastic gasket. The two ends of the main bearing core body are of 60-degree cone structures and are matched with the conical face of the main bearing core body through a conical cushion block and an indexing cushion block assembly formed by splicing an indexing thread locking block and an indexing inclined cushion block, clamping of the ultra-thin-wall-thickness interlayer is achieved under the pretightening force effect of an elastic gasket, and finally locking and fixing are conducted through a locking nut. All the components are made of continuous SiC fiber reinforced titanium-based composite materials through hot press molding. Through the synergistic effect of conical surface matching and elastic pre-tightening, uniform distribution of stress at the connecting position is achieved, damage to an ultrathin interlayer is effectively avoided, and the connecting piece has the advantages of being resistant to high temperature (900-1000 DEG C), high in strength (sigma b is larger than or equal to 700 MPa-900 MPa), capable of preventing falling, locking, high in vibration resistance and the like, and is particularly suitable for connection of a high-speed aircraft composite material interlayer structure.

The application discloses a fan wing trailing edge adjusting device for flight and relates to the technical field of aircrafts, which comprises a trailing edge angle adjusting mechanism, a fairing angle adjusting mechanism, a circular end plate and a fairing plate, the trailing edge angle adjusting mechanism comprises a steering engine, a first connecting rod mechanism, a trailing edge plate and a rotating shaft, the steering engine is fixedly installed on the circular end plate, the steering engine controls the trailing edge plate to move through the first connecting rod mechanism, the rotating shaft is arranged on one side of the trailing edge plate, and the trailing edge plate is hinged to the fairing plate through the rotating shaft, the size of the trailing edge angle and the fairing angle is freely adjusted, the aerodynamic force of the fan wing aircraft changes, the lift and thrust distribution is improved, the aerodynamic effect is enhanced, the energy consumption is reduced, and the flight efficiency is improved. The fan wing trailing edge adjusting device is applied to a tandem fan wing, the size of the trailing edge angle and the fairing angle is changed, the positive coupling gain between the two wings is realized, different flight conditions are met, various flight requirements are flexibly coped with, and the load performance and the endurance are optimized.

A lightweight, heat-insulating, load-bearing composite structure for wings and rudders. The present invention belongs to the field of aerospace and marine manufacturing technology. It solves the problem that existing heat-insulating, load-bearing pipe structures cannot simultaneously meet the requirements of heat protection, load bearing, and heat-free short-circuit connection. Structure: The middle part of the outer circumference of the inner metal tube is covered with a core layer tube, the middle part of the outer circumference of the core layer tube is covered with a ring layer tube, and the middle part of the outer circumference of the ring layer tube is covered with an outer metal tube; a load-bearing base and a heat-insulating interlayer are provided between the core layer tube and the ring layer tube. The present invention is a lightweight, heat-insulating, load-bearing composite structure for wings and rudders.

This invention relates to a self-driven active cooling system for aircraft control surfaces, belonging to the field of aircraft thermal protection technology. It includes: a pressurization pipeline and a jet pipeline disposed on the control surface; the pressurization pipeline contains a phase-change working fluid, and the jet pipeline has nozzles facing the high heat flux region of the control surface; a vapor collection chamber and a liquid storage chamber disposed on the fuselage, separated by a partition; an elastic reset member disposed in the liquid storage chamber, acting on the partition; the vapor collection chamber is connected to the outlet of the pressurization pipeline and the inlet of the jet pipeline; the liquid storage chamber is connected to the inlet of the pressurization pipeline through a return pipeline. The system of this invention can achieve self-driven, high-efficiency heat exchange.

The application belongs to the field of additive manufacturing design and manufacturing integration, and particularly relates to a 4D printed heat pipe type leading edge structure. The structure comprises a skin, which comprises an outer skin and an inner skin, the inner skin is arranged on the inner side of the outer skin, the outer skin and the inner skin have a containing cavity therebetween, the containing cavity contains liquid working medium, the inner surface of the outer skin and the inner surface of the inner skin are both provided with fine dot array structures; a plurality of partitions are arranged between the outer skin and the inner skin, and are used for separating the containing cavity, the partitions are provided with through holes, and the through holes are used for the flow of liquid working medium in adjacent containing cavities; a body connecting structure is arranged on the skin, and is used for connecting the skin and a body structure. The application can solve the problem of thermal protection of the leading edge structure of a hypersonic aircraft, and has the advantages of compact structure, high practicability and high reliability.

A water vehicle including a fluid system is described. An example water vehicle includes a fluid system, a hull, a sail, and a compressor. The hull has a bow and a stern. The sail has a first end, a second end disposed on the hull, a longitudinal axis, a central axis, a first side, a second side opposite the first side, and defines a spray opening, a suction opening, and a channel. The longitudinal axis extends from the first end to the second end. The central axis is arranged orthogonal to the longitudinal axis and is located between the first side portion and the second side portion. The suction opening is disposed on the second side and between the injection opening and the first side. A passage extends from the suction opening to the ejection opening such that fluid can travel into the suction opening and out of the ejection opening. A compressor is disposed within the passage.