Abstract: A thermal management system for an aircraft fuel system includes a plurality of fuel pressure sources. A spool valve is fluidly connected to the pressure sources and is movable between a plurality of positions in response to the pressure sources. An oil cooler is fluidly connected to the spool valve with fuel flowing through the spool valve to the oil cooler in response to the spool valve being in one of the positions. The fuel tank is fluidly connected to the spool valve with fuel flowing to the spool valve to the fuel tank in response to the spool valve being in another of the positions. The spool valve moves to a closed fuel tank output position response to the pressure sources. The spool valve includes a housing having a bore with the plurality of pressure inputs and oil cooler and fuel tank outputs fluidly connected to the bore. A spool is disposed within the bore and is movable axially relative thereto between oil cooler, open fuel tank and the closed fuel tank outward positions.

Abstract: A passive cooling system for an auxiliary power unit (APU) installation on an aircraft is provided. The system is for an auxiliary power unit having at least a compressor portion of a gas turbine engine and an oil cooler contained separately within a nacelle. The system includes the auxiliary power unit housed within the nacelle of the aircraft, an engine exhaust opening defined in the aft portion of the nacelle and communicating with the gas turbine engine, at least a first air inlet duct communicating with a second opening defined in said nacelle and with said compressor portion and the oil cooler is located within a second duct communicating with an opening other than the engine exhaust opening of said nacelle and with the engine exhaust opening. Exterior cooling air and engine exhaust ejected through said engine exhaust opening entrain cooling air through said second duct to said oil cooler, and thus provide engine oil cooling. An exhaust eductor is also provided.

Abstract: A heat exchanger system and method for heating selected areas of the wings of an aircraft. The heat exchanger apparatus involves the use of a heat exchanger and thermal contact with the fluid of a gear box associated with a geared fan powerplant used on an aircraft. The heat exchanger is in fluid communication with at least one, and more preferably a plurality, of conduits which circulate fluid heated by the fluid of the gearbox throughout the selected areas of the wings of the aircraft. This serves to selectively heat various areas of the wings to help prevent icing of the leading edge of each of the wings, as well as to increase the region of laminar flow over the wings, and thus decrease those areas where turbulent flow occurs.

Abstract: A passive cooling system for an auxiliary power unit (APU) installation on an aircraft is provided. The system is for an auxiliary power unit having at least a compressor portion of a gas turbine engine and an oil cooler contained separately within a nacelle. The system includes the auxiliary power unit housed within the nacelle of the aircraft, an engine exhaust opening defined in the aft portion of the nacelle and communicating with the gas turbine engine, at least a first air inlet duct communicating with a second opening defined in said nacelle and with said compressor portion and the oil cooler is located within a second duct communicating with an opening other than the engine exhaust opening of said nacelle and with the engine exhaust opening. Exterior cooling air and engine exhaust ejected through said engine exhaust opening entrain cooling air through said second duct to said oil cooler, and thus provide engine oil cooling. An exhaust eductor is also provided.

Abstract: An inverted aircraft internal combustion engine, such as a two-stroke compression-ignition (diesel) engine, incorporates a lubricant (oil) and/or coolant heat exchanger mounted upon the engine below an exposed crankshaft and/or enclosed crankshaft housing and adjacent a propeller mounting location.

Abstract: A leading edge member of an airfoil of an aircraft has an outside wall defining a wing span and a deicing device for circulating hot air in the direction of the wing span of the airfoil. The device includes a hot air distribution duct extending inside the leading edge member in the direction of the span of the foil. The duct has a plurality of orifices distributed over at least part of its length. A hot air supply communicates with one end of the distribution duct. The hot air supply is connected by ducts located outside the leading edge member to at least one propulsion engine of the aircraft.

Abstract: A cooling device for an aircraft turboshaft engine includes a boundary layer suction device formed by a perforated external wall portion of the aircraft, a collector on the inside of the wall portion for receiving the air flowing in through the perforations, and a discharge device discharging the air from the collector. The cooling device is completed by a heat exchanger which is disposed in the path of the air flow through the collector, a pump for pumping fluid to be cooled through the heat exchanger from a circuit of the engine, and a regulator regulating the fluid flow through the exchanger to control the temperature of the fluid. The fluid which is cooled may be from the lubrication circuit, the fuel circuit, or an air circuit of the engine, and the boundary layer suction device is preferably disposed in the pod of the engine.

Abstract: A liquid coolant heat exchange system for use in semimonocoque aircraft includes an arcuate planar heat sink fixed along a radius of curvature R.sub.1 by forming members, a flexible arcuate planar spreader plate having a radius R.sub.2, such that R.sub.1 >R.sub.2, a heat exchange tube for transferring heat from the liquid coolant to the heat sink and means for attaching the spreader plate to the forming members to hold the spreader plate in contact with the heat sink.

Abstract: A liquid coolant heat exchange system for use in semimonocoque aircraft includes an arcuate planar heat sink fixed along a radius of curvature R.sub.1 by forming members, a flexible arcuate planar spreader plate having a radius R.sub.2, such that R.sub.1 >R.sub.2, a heat exchange tube for transferring heat from the liquid coolant to the heat sink and means for attaching the spreader plate to the forming members to hold the spreader plate in contact with the heat sink.

Abstract: A three axis stabilized geostationary satellite comprises a structure having walls that face North and South when the satellite is on station and carrying at a storage battery of large capacity. The storage battery is constituted by cells mounted on a chassis having a base plate oriented North-South and mechanically secured to sides that face North and South when the satellite is on station. The sides carry thermal radiators that are thermally connected to each other and to the cells. Typically sides are in alignment with North and South faces of the structure and extend away from the Earth.

Abstract: Cylindrical battery cell vessels are coupled together in plural sets by pairs of half-shell sleeves. The sleeves conduct heat preferentially in an axial direction. Each sleeve set is mounted onto a heat rejection plate for direct radiation to space. Each satellite of a set of spacecraft has a North and/or South-facing region for mounting the heat rejection plates of the requisite number of cells. Each plate may be fitted with an optical solar reflector (OSR) for reflecting insolation, and for radiating heat energy.

Abstract: An air/liquid heat exchanger having a refrigerant spray providing evaporative cooling.

Abstract: A cooling system for a hypersonic aircraft is disclosed which enables hypersonic flight using non-cryogenic fuels. The cooling system positions a primary heat exchanger at an external location on the aircraft which remains relatively cool during hypersonic flight. A working fluid is passed through the primary heat exchanger to the hot parts of a supersonic combustion ram jet engine.

Abstract: An air/liquid heat exchanger having a tube that reinforces the heat exchanger's core and header plates and also sprays a liquid onto the core for evaporative cooling.

Abstract: An air/liquid heat exchanger for aircraft having ram pressure powered evaporative cooling.

Abstract: A pusher turboprop engine is provided with an interior exhaust duct structure which directs the hot turbine gases through and out of the engine nacelle and around the rear spinner which surrounds the propeller hub. The spinner supports blade-shaped shields which protect the roots of the propeller blades from the hot exhaust gasses.

Abstract: A pusher turboprop engine is provided with an interior exhaust duct structure which directs the hot turbine gasses through and out of the engine nacelle to an annular duct mounted on the rear spinner which surrounds the propeller hub. The rotating annular duct includes blade-shaped shields which protect the roots of the propeller blades from the hot exhaust gasses and also pulls warmed cooling air through the engine nacelle thereby providing a rearwardly directed jet thrust to augment the propeller thrust.

Abstract: A helicopter is driven by means of a rotor that is connected with an internal-combustion engine. The air-cooled internal-combustion engine has a horizontally opposed construction and is arranged within walls of a cabin of the helicopter. In order to supply the internal-combustion engine with cooling air and intake air, air inlet openings are provided at the side walls of the cabin. For the discharge of the cooling air, outlet openings are arranged at the underside of the cabin.

Abstract: An unmanned aircraft is equipped with a helium-cooled Brayton cycle nuclear reactor. Heated helium gas drives turbines which in turn rotate propellers to maintain the aircraft aloft for a protracted period of time. After the helium gas expands in the turbines, it is passed through a closed loop including radiator tubes which radiate waste heat from the helium gas to space. The cooled helium gas is returned through the closed loop for repeat of the Brayton cycle.

Abstract: In order to dissipate aerodynamic heating in hypersonic aircraft, while at the same time providing structural support for the leading edge of an airfoil or the nose cone in a manner that avoids transmittal of thermal bending loads into the aircraft, a structural cooling unit is formed of a load bearing structure. The load bearing structure is of hollow construction defining a fluid flow path therethrough and has a fluid inlet at one end of the fluid flow path and a fluid outlet at the other end of the fluid flow path which comprises a tortuous path through the load bearing structure. Additionally, for purposes of forming a structural load path, the load bearing structure is formed into an elongated tubular configuration having at least one substantially continuous heat exchanging surface for aerodynamic heat dissipation.

Abstract: An ultra hypersonic aircraft having aerodynamic heat transferring pipes forming an airframe structure is shown. The heat transformer pipes include a high pressure steam generator located at the forward section thereof and a low pressure steam condensator located at the rearward section thereof. Low pressure water is circulated between the rearward section and the forward section for heat balancing of the air frame structure and for operating the room air conditioning system of an aircraft. A steam pressure reactant turbine, which drives an electric generator, is used as a boundary between the high and low pressure steam sections. The electric generator provided electrical and power supply for the infrared or arc heaters, and electric heaters recessed in the ceiling and wall panels of an acrodynamic lift-thrust generating channel for back-heating of the turbo-ram oval thrust stream in the thrust generating channel of a hypersonic aircraft.

Abstract: An aircraft having a rotatively driven blade cooperating with an electrochemical power system for operating an electric drive system. The blade provides cooling heat transfer relative to the lithium/peroxide fuel system effecting desirable separation of the lithium-hydroxide monohydrate in a suitable filter/plenum provided as a portion of the power system. In one form, the cooling is effected by disposing a heat transfer portion of the power system in the path of air flow generated by the propulsion blade of the aircraft. In another form, the cooling is effected by flowing the electrolyte through the propulsion blade. Molarity of the electrolyte is maintained by separating the precipitated lithium-hydroxide monohydrate in solid form in the filter/plenum and addition of replacement hydrogen/peroxide solution from the reservoir thereof.

Abstract: Cooling for an electronic package mounted on an aircraft engine housed in a nacelle is achieved by connecting the package to the inlet of the engine and overboard of the nacelle so that substantially continuous air flow over the electronic package will ensue from either direction depending on the varying pressure gradients adjacent the points where connections are made and the conditions exists during the flight envelope. This negates the necessity of providing positive pumping devices for effectuating the flow of the cooling air.

Abstract: In combination with an internal combustion engine and an air compressor, for use in an aircraft, the compressor providing compressed air to the engine's air inlet, there is provided a variable overboard bleed valve. The bleed valve is operated by a valve controller which controls the position of the bleed valve in response to changes in engine operating conditions. This permits maintaining optimum compressor operation over a wide range of engine operating conditions without exceeding the critical pressure of the compressor, so as to avoid surge. The controller can be preferably linked to the propeller governor and the throttle control for the aircraft, so as to move in response to movement of the governor and throttle control.

Abstract: In combination with an internal combustion engine and an air compressor, for use in an aircraft, the compressor providing compressed air to the engine's air inlet, there is provided a variable overboard bleed valve. The bleed valve is operated by a valve controller which controls the position of the bleed valve in response to changes in engine operating conditions. This permits maintaining optimum compressor operation over a wide range of engine operating conditions without exceeding the critical pressure of the compressor, so as to avoid surge. The controller can be preferably linked to the propeller governor and the throttle control for the aircraft, so as to move in response to combined movement of the governor and throttle control. The combined movement can be obtained using a pair of cables connected to a pivoting plate connected to the rotating control shaft of a butterfly valve.

Abstract: An air inlet control inset in the skin or outer wall surface of an aircraft or the like. The control device has an open position in which it is able to capture ram air flowing over the wall surface and which at the same time defines an open flow area through which lower pressure ambient air can be inducted when ram air is unavailable. The air inlet control includes a shutter opening into ram air flow, and an actuating mechanism therefor. In one disclosed form of the invention, the shutter is aerodynamically biased to reduce actuating forces required to close the shutter against ram air pressure.

Abstract: A structure is provided for cooling the tail rotor gearbox of a single rotor helicopter having an anti-torque tail rotor mounted to a vertical stabilizer. The gearbox is mounted within the vertical stabilizer inside an air passage passing through the stabilizer. An air inlet is located in the vertical stabilizer on the side thereof opposite the tail rotor. An exhaust orifice is positioned in the vertical stabilizer on the same side of the stabilizer as the tail rotor. In operation the pressure differential over the vertical stabilizer caused by the rotation of the rotor causes ambient air to enter the inlet, pass through the passage, over the gearbox, and exit from the exit orifice thereby cooling the gearbox.

Abstract: An airplane with an electric motor and rechargable battery pack are disclosed herein. The electric motor is of the high performance variety (i.e., it has a very high power rating for its size and weight). Means are provided for producing a high velocity blast of cooling air directly onto the rotating armeture for rapidly dissipating the waste heat generated inside of the motor. Means are also provided for obtaining a large power output from the battery pack and for very rapidly recharging the battery pack. A remote control system is included for controlling the operation of the airplane from a remote location.