Abstract: This invention relates to a method and an apparatus for the cooling of the casing (1) of the turbines (2, 3) of jet engines, in which cooling air is diverted from a bypass flow and supplied to the outer side of the casing via an inlet duct (8) provided with a shut-off element (5). According to the present invention, the cooling air is supplied to a first chamber (6) in which it is divided by volume. One portion of the cooling air is issued to the casing (1) via orifice holes (10), while another portion is ducted via several tubes (9) to a second chamber (7) which annularly encloses the casing (1) in the area of a low-pressure turbine (3).

Abstract: The cooling of the exhaust nozzle of a fan jet engine powering aircraft is enhanced by providing an extension cooling concept that includes an extension piece extending axially downstream of the nozzles variable throat that utilizes fan air to cool the forward portion of the divergent nozzle surfaces adjacent to the gas path of the engine and ram air to cool the surfaces downstream of the ejector. A variable vane varies the area of the ejector slot as a function of Aj for compensating for flow losses at off-design points of the nozzle positions during the flight envelope of the aircraft so as to avoid overtemperature of the components.

Abstract: An exhaust induced ejector nozzle (20) includes a convergent portion (30) and a divergent portion (34) disposed downstream from the convergent portion (30). The ejector nozzle (20) also includes an ejector (40) having an outlet (46) disposed within the divergent portion (34). The ejector (40) is operable to provide entrainment of a nacelle airflow (18) from a nacelle airflow area (44) into the divergent portion (34). The ejector nozzle (20) further includes a convergent ejector (60) extending between the convergent portion (30) and the nacelle airflow area (44). The convergent ejector (60) is operable to draw a portion of an exhaust airflow (42) into the nacelle airflow area (44) to increase a flow rate of the nacelle airflow (18).

Abstract: A device comprises a bellows with protective rings arranged between corrugations of the bellows. Support rings are connected in a sealed manner to a gas line and a combustion chamber. A gimbal ring is positioned outside the bellows and via hinges is connected to structural brackets with support rings. A shield is mounted inside the bellows, the shield consisting of two cylindrical envelopes telescopically inserted one in another with a gap. The cylindrical envelopes are fixed in cantilever to support rings, forming a chamber which via channels for feeding a cooling working medium made in the support rings is connected to a main line for feeding the cooling working medium, and via the gap between the envelopes—to the cavity of the bellows unit. A housing is arranged outside the protective rings and adjoining them, and is made in the form of a metallic cylindrical helix, the ends of which are connected to the support rings (FIG. 2).

Abstract: The cooling of components (26) of a vehicle (10) carrying a turbine engine (12) and a fuel supply (24) for the engine (12) without the excessive consumption of bleed air is achieved in an apparatus which includes a fuel recirculating system (34), a first heat exchanger (60) for rejecting heat from selected components (26) of the vehicle (10) to fuel flowing through the fuel recirculating system (34), and a cooling device (32) including a second heat exchanger (100, 202) for rejecting heat from the fuel to the cooling device (32) as the fuel flows through the fuel recirculating system (34).

Abstract: The wall downstream of the slot of an ejector forms a divergent angle relative to the upstream wall adjacent to the slot so that the flow discharging from the slot expands and avoids having the primary air from pinching the secondary air and adversely affecting the cooling film issuing from the slot. When multiple slots in the ejector are utilized the downstream wall in relation to the upstream wall is similarly configured to obtain local multi-expansions.

Abstract: The housing is heated or cooled to adjust its diameter and, in particular, the play between the tips of the rotor blades and the housing. The apparatus consists of networks of tubes (1) covering complementary portions of the circumference of the housing and consisting of a pair of distributors (3) and parallel tubes (2) between said distributors, said tubes (2) being alternately connected to one or other of the distributors and provided with apertures to blow the gas onto the housing. The counter-flow circulation in the tubes (2) makes it possible to supply both gas that has been greatly heated by traveling a considerable distance through the tubes (2) and cooler gas that has traveled less far in other tubes (2) to any point on the housing, thereby balancing out the heat applied to every portion of the housing.

Abstract: A turbine engine (20) having bypass and core passages (30 and 40), a high pressure compressor (80) for compressing air in the core passage and an intercooler (120, 220) for cooling core air (42) in the core passage prior to discharge by the high pressure compressor. The intercooler includes a plurality of heat exchange elements (140) positioned in the bypass passage through which all or a portion of the core air is transported. Heat in the core air is transferred via the heat exchange elements to the relatively cool bypass air, thereby cooling the core air and heating the bypass air. The turbine engine may have a counterflow intercooler (120), a radial flow intercooler 220 or an intercooler of other configuration. The turbine engine may have a high pressure compressor with two sections (82 and 86), with core air transported to and from the intercooler between sections, or a single section high pressure compressor, with core air transported to and from the intercooler upstream of the high pressure compressor.

Abstract: An IR Suppressor (20) operative to suppress the infrared signature radiated from the high-temperature exhaust of an engine (12) and comprising a primary exhaust manifold (22), first and second mixing ducts (24a, 24b), and a secondary flow shroud (26). More specifically, the primary exhaust manifold (22) includes an elongate duct (28) adapted for receiving a primary flow of the high-temperature engine exhaust from the engine and at least two high aspect ratio nozzles (30a, 30b) integrated in combination with the elongate duct (28). The high aspect ratio nozzles (30a, 30b) are positioned so as to define at least two outlets at the same axial station along the length of the elongate duct (28). The first and second mixing ducts (24a, 24b) are disposed over the high aspect ratio nozzles (30a, 30b) so as to form a mixer/ejector defining at least two ejector inlets (44).

Abstract: A flow diverting mechanism for the twin duct offtakes or coannular offtake for a turbo fan engine powering aircraft with either thrust reversing or short takeoff and vertical landing or both capabilities includes a rotary drum mounted between the turbine and afterburner sections and includes ports complementing the inlets in the offtakes to divert the core stream through the offtakes and passages to proportion the fan discharge air utilized for cooling purposes in the engine to cool the liner and walls of the offtakes and the engine components located downstream of the rotary drum while assuring separation of the core stream flow and fan air flow.

Abstract: An inlet-outlet module 108 having both an inlet 58 and an outlet 62 for a heat exchanger system 52 is disclosed. Various construction details which improve the stiffness to weight ratio and the aerodynamic smoothness are developed. In one embodiment, the module 108 is formed of a one piece casting having a lobed mixer 94 at the outlet.

Abstract: A thrust chamber for an expander cycle rocket engine having a predetermined quantity of tubes concentric about an axis and radially inward of a structural jacket, each tube having first and second substantially planar sidewalls which diverge at an angle such that immediately adjacent tubes are in spaced relation to each other at the radially inward portion of each tube.

Abstract: A flow diverting mechanism for the twin duct offtakes or coannular offtake for a turbo fan engine powering aircraft with either thrust reversing or short takeoff and vertical landing or both capabilities includes a rotary drum mounted between the turbine and afterburner sections and includes ports complementing the inlets in the offtakes to divert the core stream through the offtakes and passages to proportion the fan discharge air utilized for cooling purposes in the engine to cool the liner and walls of the offtakes and the engine components located downstream of the rotary drum while assuring separation of the core stream flow and fan air flow.

Abstract: A turbofan engine augmenter includes an exhaust casing and a combustion liner therein which define therebetween a cooling duct for receiving bypass air. A diffusion liner adjoins the combustion liner and defines with the casing an outer inlet for receiving the bypass air, and a radially inner inlet for receiving core gases. A ring flameholder includes an aft facing radial wall adjoining the diffusion liner, and an integral axial wall adjoining the combustion liner. A plurality of circumferentially spaced apart radial flameholders extend inwardly from the casing and through the diffusion liner forward of the ring flameholder, and each includes a first fuel spraybar for injecting fuel into the core gases. A mixer defined by a plurality of injector chutes extends radially inwardly through the diffusion liner, with the chutes being spaced between respective pairs of the radial flameholders.

Abstract: A turbojet engine includes a compressor and turbine, with a combustor disposed therebetween for generating exhaust gas. An afterburner includes an annular casing and liner spaced radially inwardly therefrom to define a cooling duct therebetween. A recoup duct is disposed in flow communication with the compressor for receiving a portion of compressed air therefrom as recoup air. A plenum is disposed at an upstream end of the afterburner casing, and includes an inlet disposed in flow communication with the recoup duct for receiving the recoup air, and an outlet disposed in flow communication with the afterburner cooling duct for discharging the recoup air therein.

Abstract: A bypass air valve is provided comprising a liner, a strap, and apparatus for selectively actuating the valve. The liner includes an inner surface, an outer surface, a plurality of first regions, and a plurality of second regions. Each first region includes a plurality of first apertures. The strap includes a plurality of openings and a plurality of third regions. The third regions include a plurality of second apertures. The valve may be selectively actuated into an open position where the first regions are in communication with the openings, and the third regions are in communication with the second regions. The valve may be selectively actuated into a closed position where the first regions are in communication with the third regions, and the second regions are in communication with the openings.

Abstract: The exhaust in a gas turbine engine is lined with stacks of hollow cooling blocks made of a composite material. Air is pumped into each stack and escapes into the exhaust flow from a slot between the blocks in each stack. In addition, the blocks contain passages so that some of the airflow escapes to a space between the stacks. The blocks are attached using special fasteners that does not create mechanical stresses on the composite material.

Abstract: A combustor, for a gas turbine engine, employing regenerative combustor cooling. The combustion gas flow direction extends generally longitudinally aft of the combustor fuel nozzle. A coolant flowpath between the combustor casing and the combustor liner has: 1) a longitudinally aft inlet in fluid communication with a source of compressor-derived cooling air, of lower temperature and higher pressure than diffused air from the combustor diffuser; and 2) a longitudinally forward outlet in fluid communication with the combustor fuel nozzle for "spent" cooling air to be used for combustion.

Abstract: A nozzle cooling system for selectively cooling longitudinally extending and circumferentially adjacent divergent exhaust flow confining elements bounding a hot exhaust gas flowpath in a divergent section of an aircraft gas turbine engine exhaust nozzle and providing pivoting capability for divergent flaps and seals in an axisymmetric vectoring nozzle. The apparatus includes axially adjacent forward and aft sections of at least one of the exhaust flow confining elements (typically referred to as flaps and seals), the adjacent forward and aft sections have forward and aft interior hot surfaces respectively, and mounting apparatus for mounting the aft section to the forward section in one of at least two positions.

Abstract: In a premixing burner for the operation of an internal combustion engine, a combustion chamber of a gas turbine group or a firing system, which burner can be used for the combustion of fuels with very different properties, a diffuser (4) is arranged after a swirl generator (1), a convergent part (2) and a throat (3) for the supply of the fuel (9). The convergent part (2), the throat (3) and the diffuser (4) have a rotationally symmetrical configuration and the variation in cross section is adapted in such a way that no flow separation is possible at the wall. Gaseous fuels having a high flame speed are introduced counter to the swirl generated by the swirl generator (1). Gaseous fuels having a low flame speed are introduced in the same direction as the swirl generated by the swirl generator (1).

Abstract: To reduce a noise and to cool a nozzle in a supersonic jet propelling engine, a sound absorbing member made of a porous plate (4) and a honeycomb structure (2) or a porous plate (4) and a torus core (3) is provided between a nozzle plate (5) and a liner (1), a cooling air (7) is caused to flow along an inner surface of the nozzle plate (5) and to impinge against the liner (1) through holes (10) of the porous plate (4). Furthermore, a number of holes (11) which are slanted on a rear side are formed in the liner (1) so that the cooling air (9) are injected through the holes (11).

Abstract: In a cooled wall part having a plurality of separate convectively cooled longitudinally cooling ducts (2) running near the inner wall (1) and parallel thereto, adjacent longitudinal cooling ducts (2) being connected to one another in each case via intermediate ribs (3), there is provided at the downstream end of the longitudinal cooling ducts (2) a deflecting device (4) which is connected to at least one backflow cooling duct (6) which is arranged near the outer wall (5) in the wall part and from which a plurality of tubelets (7) extending to the inner wall (1) of the cooled wall part and arranged in the intermediate ribs (3) branch off. By means of this wall part, the cooling medium can be put to multiple use for cooling (convective, effusion, film cooling).

Abstract: A diverging rocket engine nozzle has a system to allow ambient gases to flow from the outside to the inside of the nozzle's chamber. The system includes a plurality of coolant conduits adjacent the inside surface of the nozzle. The conduits are spaced apart to provide a plurality of vent openings which extends from the inside to the outside of the nozzle's chamber. A plurality of valves controls gas flow through the vent passages when the pressure outside of the nozzle is greater than the pressure directly adjacent the inside surface of the nozzle. The coolant conduits cool and shield the valves at their points of attachment to the outside of the nozzle.

Abstract: A cooled multi-wall liner is provided having an outer wall, an inner wall and a porous midsection disposed therebetween, attached to both walls. The outer wall includes a plurality of first apertures. The inner wall includes a plurality of sections separated by gaps. Fluid on the outer wall side of the liner at a pressure greater than fluid on the inner wall side of the liner may enter the liner through the first apertures, be diffused within the porous midsection, and exit the liner through the gaps between the sections of the inner wall.

Abstract: The nacelle of the aircraft is purged and the sidewall of the vector nozzles of the gas turbine engine in the nacelle is cooled by utilizing the engine fan air to drive ejector pumps mounted in series flow relationship, a rectangular mixing zone and a diffuser with a rectangular cross section and a step in the exhaust nozzle gaspath. This results in minimizing the amount of fan air for driving the system and enhancing engine thrust and enhanced sidewall cooling because the temperature of the mixed air is lower than heretofore known systems.

Abstract: This disclosure relates to a system for purging of the nacelle of an aircraft and for cooling the edge pieces of the divergent flap of a gas turbine engine's two-dimensional exhaust system by use of long tubular air pump with axisymmetric supersonic nozzles, a rectangular mixing chamber for pumping nacelle air with fan discharge air and judiciously flowing the air in separate passageways formed in the double clamshell convergent flap of the exhaust nozzle. This reduces the overall aircraft and weapon system weight, with only a minimum increase in the engine's nozzle weight.

Abstract: A method and apparatus provide cooling air from a compressor to hot components in a gas turbine engine. Compressed air discharged from an exit of the compressor is stratified into a main flow channeled to a combustor, and a separate bleed flow channeled to a heat exchanger. The bleed flow has an above average pressure whereas the main flow has a below average pressure. The bleed flow is cooled in the heat exchanger and channeled to hot components of the engine for cooling thereof.

Abstract: The apparatus is a cooling structure for reaction engine throats. The constriction in the throat is cooled by a group of heat pipes which radiate outward from the constriction to a larger diameter perimeter surface where the heat is dissipated. The entire structure can be constructed by embedding pretested heat pipes around a base structure with plasma sprayed metal.

Abstract: Cooling of the hot surfaces of the exhaust nozzle of a fan jet engine and purging of the nacelle is attained by a booster pump mounted in exhaust nozzle structure and the booster pump having an ejector driven by localized fan air existing in the engine cooling path extracted from the fan of the engine for pumping nacelle air and mixing the egressing fluid in a mixing chamber and diffuser disposed in fluid communication with booster pump which is then dumped into a plenum within the exhaust nozzle structure for feeding film cooling holes formed in the hot surfaces for laying a blanket of cooling air along the gas path surface of the exhaust nozzle.

Abstract: An impingement cooling arrangement which avoids thermal cycling fatigue failures in the cooled member and cooling baffle member while retaining permanent rigid attachment of the cooling baffle member and cooled members at their periphery. The baffle includes an array of stress relieving smoothly curved in cross section furrows or grooves dispersed over the baffle surface and provides for intentional heat conduction to the furrow valley portions as a stress relieving mechanism. Heat conduction to the furrow valley portions is aided by the force of pressurized impingement cooling air at the baffle surface.

Abstract: There is disclosed an anode for an electrothermal arcjet thruster. The anode includes a divergent nozzle having, in tandem, a recombination portion and an expansion portion. The expansion portion has a greater rate of divergence than the recombination portion. A regeneration channel containing flowing propellent gas extends internally within the anode adjacent the recombination portion. The combination of the biangle nozzle and the regeneration channel produces an arcjet thruster having a markedly increased thrust efficiency.

Abstract: The present invention provides a cooling apparatus with two spaced apart walls, one of which is a coolable wall, having a cooling flowpath between them wherein the cooling flowpath is converging in the direction of the cooling flow. In one embodiment the coolable wall is an afterburner cooling liner that includes a film cooling means on the hot side of the liner. Another embodiment provides a corrugated or wavy wall afterburner cooling liner wherein each wave includes a forward facing surface, from crest to trough of the wave in the direction of the hot gas flow, and an aft facing surface, from trough to crest. A high density pattern of multi-hole film cooling holes is disposed on the hotter of the two surfaces and a lower density pattern on the cooler one.

Abstract: In a rocket engine nozzle with a selectively smaller outlet cross-section, the nozzle comprising a converging portion which receives the gases produced in a combustion chamber, a nozzle throat of small cross-section and a diverging part connected to the nozzle throat and terminating at its downstream portion in a gas jet outlet cross-section defining a high cross-section ratio, a releasable annular obstacle is connected to the downstream end of the diverging part so as to define at the outlet of the diverging part a zone of selectively smaller cross-section which ensures stability of the separation of the flow of hot gases relative to the wall of the diverging part during a first phase of flight.

Abstract: The present invention provides a cooling apparatus with two spaced apart walls, one of which is a coolable wall, having a cooling flowpath between them wherein the cooling flowpath is converging in the direction of the cooling flow. In one embodiment the coolable wall is an afterburner cooling liner that includes a film cooling means on the hot side of the liner. Another embodiment provides a corrugated or wavy wall afterburner cooling liner wherein each wave includes a forward facing surface, from crest to trough of the wave in the direction of the hot gas flow, and an aft facing surface, from trough to crest. A high density pattern of multi-hole film cooling holes is disposed on the hotter of the two surfaces and a lower density pattern on the cooler one.

Abstract: The apparatus is a cooling structure for reaction engine throats. The constriction in the throat is cooled by a group of heat pipes which radiate outward from the constriction to a larger diameter perimeter surface where the heat is dissipated. The entire structure can be constructed by embedding pretested heat pipes around a base structure with plasma sprayed metal.

Abstract: An ejecting exhaust nozzle in a gas turbine engine includes a convergent section, a divergent section, extending downstream from the convergent section, and an ejector for ejecting an airflow into the core airflow is fed by a boost pump. The boost pump includes a flow mixer with a boost plane wherein a fan airflow is injected into a ram airflow and then mixed in a mixing section extending downstream from the boost plane. Varying the fan airflow to boost the ram airflow to control a boosted ram airflow pressure allows ejector operation throughout the flight envelope of the gas turbine engine.

Abstract: A planar top plate (32) has film cooling slots (34) therein. A rigid honeycomb structure (38) sealingly joins this top plate with bottom plate (36). Cooling air distributed by openings (50) impinges on the top plate and passes through metering openings (48) to the slots. A flat wall surface with reduced aerodynamic losses is maintained and the uniform support of the top plate avoids local distortion which would crack any heat resistant coating applied thereto.

Abstract: The hot core engine compartment of a turbojet engine is cooled by the metered flow of cooling air exhausted through a laminar flow nacelle system. This arrangement reduces the need for extracting bypass ram air from the engine performance cycle for cooling purposes and thereby improves specific fuel consumption and engine efficiency. A turbocompressor pump driven by bleed air from the core engine compressor draws cooling air through the laminar flow nacelle system and pumps the air into a manifold surrounding the core engine.

Abstract: An exhaust nozzle has variable convergent flaps, ejector flaps, and short divergent flaps spaced inwardly from the ejector flaps, cooling air is introduced through the opening between the divergent and ejector flaps. When the nozzle is vectored, the size of the opening is increased in the region of the nozzle experiencing higher static pressure.

Abstract: A cooling system for an axial flow compressor of a gas turbine engine is disclosed having a device to tap a portion of air from the primary air flow flowing through the axial compressor and a control device associated with the tap to control the amount of air tapped from the primary air flow. The system also includes a cooling device to cool the portion of air tapped from the primary air flow, which cooled air is directed into a cooling air chamber. The cooling air chamber is located between the last rotor stage of the compressor and a cone attached to the last stage and extending downstream from the rotor wheel. Nozzles associated with the cooling air chamber direct cooling air from the chamber onto the rotating rotor wheel such that a portion of the cooling air passes into the peripheral groove defined by the rotor wheel between the wheel and the blade root, while another portion of the cooling air from the nozzles is directed onto the surface of the cone in order to cool the blades and the cone.

Abstract: A fan-jet gas turbine engine heat exchanger is cooled by fan air taken from the fan duct at a relatively high pressure region of the fan nozzle upstream of the fan nozzle's throat and exhausted into a relatively low pressure region of the fan nozzle upstream of the throat. One embodiment provides a cooling air duct with a forward facing scoop at the duct's inlet for capturing cooling air and providing greater dynamic pressure recovery. A diffuser is located upstream of the heat exchanger and a nozzle is located downstream of the heat exchanger with a nozzle exit that forms the cooling duct outlet and an aft facing scoop over the outlet.

Abstract: A cooled aerofoil blade for a gas turbine engine has a hollow trailing edge region in which bridging members interconnect the convex and concave flanks of the blade. The bridging members are each provided with angled cooling passages which are in operation exposed to a cooling air flow and direct that flow onto the internal surfaces of the convex and concave flanks so as to provide impingement cooling thereof.

Abstract: A cooling air transferring apparatus is provided particularly for transferring cooling air from hollow convergent flaps and seals to hollow divergent flaps and seals respectively in aircraft gas turbine engine variable nozzles. The preferred embodiment includes two semi-circular tubes, the first partially received within the second and both having a common center and radius of curvature so as to slide about the pivot point of the variable nozzle's pivotable convergent and divergent flaps and seals. The preferred embodiment further includes an entrainment opening on the second tube through which part of the first tube is received. The entrainment opening induces cooling air from the nozzle bay into the cooling air transfer tube to help cool the divergent flaps and seals.

Abstract: A heat exchange apparatus to heat hydrocarbon liquid fuel that is particularly useful for a combustion system and method of operation that by fuel in a heat exchanger and avoids undesirable thermal fuel deposits that would otherwise reduce the effectiveness of the heat exchanger. The invention provides a means to mix a relatively large amount of non-condensable gas with the fuel prior to gasification heating thus reducing the residence time of the fuel adjacent to the hot surface. The non-condensable gas may be air, a cryogenic such as methane, or a chemical fraction of the gasified fuel which by itself does of promote the formation of significant amounts of gums or coke deposits.

Abstract: An endothermic fuel system for cooling a heat source, which may be on a high speed vehicle, has an endothermic fuel decomposition catalyst which is capable of endothermically decomposing an endothermic fuel. The system also includes means for transferring thermal energy from the heat source to the catalyst in order to heat the catalyst to a temperature sufficient to endothermically decompose at least a portion of the endothermic fuel and to cool the heat source and means for contacting the heated catalyst with the endothermic fuel stream to cause the endothermic fuel stream to absorb the thermal energy and endothermically decompose into reaction products.

Abstract: The present invention comprises at least one active cooling apparatus incorporated into the nozzle liner of sidewall of an afterburner. The liner has a plurality of cooling air holes located under the apparatus. Within the liner is a supply of cooling air under high pressure. Over the holes in the liner is flexibly attached the active cooling apparatus which is composed of an inner manifold having a bottom layer with a plurality of receiving holes therein which are offset from the supply holes of the liner and a top layer also having a plurality of vent holes therein which are further offset from the holes in the bottom layer. Attached over the inner manifold is a hot cover layer which is located a small distance above the inner manifold top layer. The hot cover layer also has a plurality of flow path holes therein. A hot gas stream flowing over the hot cover layer bleeds into the inner manifold.

Abstract: A heat exchange arrangement in an annular fan duct of a gas turbine engine includes a conduit routing hot bleed air flow, a heat exchanger disposed within the fan duct and connected with the conduit such that hot air is passed through the interior of the heat exchanger, and at least one and preferably a pair of partition members disposed within the fan duct for temporarily diverting a small portion of the main air flow from the fan duct and over the exterior of the heat exchanger in heat transfer relationship with the hot air passing through the interior of the heat exchanger. The small air flow portion rejoins the main air flow through the fan duct and is used to generate engine thrust after cooling the hot bleed air flow carried by the conduit. The partition members disposed adjacent to opposite sides of the heat exchanger and a support member define therewith an air diffusion passage through which the small fan duct air flow portion is diverted temporarily from the main fan duct air flow.

Abstract: A gas turbine engine is provided with an annular array of rotor blades which are surrounded by a shroud ring. The shroud ring includes a number of adjacent heat pipes which are of the variable conductance type. The heat pipes each have a portion which is externally of the shroud ring which portion is in turn divided into two parts. The first part is exposed to a flow of cooling air while the second part contains a buffer which operationally varies the conductance of the heat pipe, in accordance with the temperature of air which is tapped from the high pressure compressor of the engine. This serves to ensure that the shroud ring remains substantially isothermal and provides control over the clearance between the tips of the rotary blades and the shroud ring through thermal control of the shroud ring.

Abstract: An apparatus for cooling components of an airplane engine at least partially operated by a cryogenically stored propellant via cooling air includes an air supply line for supplying the cooling air from an outside environment. A heat exchanger exchanges the cooling air with the propellant to precool the cooling air and includes a condenser having an in-flow and an out-flow side connected to the air supply line for precooling the cooling air. A pump is coupled to a cooling air line for conveying the precooled air in a liquid or vapor state to the components to be cooled wherein the cryogenically stored propellant acts upon the condenser. A propellant tank is connected to the in-flow side of the condenser and stores the cryogenically stored propellant, wherein the out-flow side of the condenser is connected to at least one combustion system of the airplane engine.

Abstract: Each convergent flap and seal has a cantilevered extension deflector. This deflector reduces the throat area at minimum throat area, and seals against a divergent flap or seal in the maximum throat area position.Cooling ambient air is induced through openings in the upstream portion of each divergent seal at intermediate throat areas. At minimum throat area the divergent flaps block these openings. At maximum throat area the deflectors block the cooling airflow path.