United Technologies Corporation

Alliant Techsystems Inc.

Orbital Atk, Inc.

Daniel P Guinan

Christopher D Gettinger

Henry K Webster

David E Wendt

James B Philpott

Alan Drake

Gary M Perkins

Donald R Miehe

Thomas B Fortin

Mark A Sillence

Sergio Rinella

Stephen A Beckel

Dean Andreadis

Revi K Nigam

Costante Salvador

Dennis J Nemecek

David J Huber

A helical cross flow pulse detonation engine.

Helical cross flow (HCF) pulse detonation engine

Methods and apparatus of an improved cooling structure having cooling passages are provided. A method includes forming pilot holes with an electrical discharge machine (EDM) drill near a heated flow surface. The pilot holes are then shaped with a wire EDM into cooling passages having a desired shape.

Method of manufacturing heat exchanger cooling passages in aero propulsion structure

In accordance with a first aspect of the present invention, an aircraft having a scramjet engine is provided. The scramjet engine includes a flowpath which varies in three dimensions. In accordance with a second aspect of the present invention, a method for launching an aircraft having a scramjet engine with a nozzle section and a solid rocket booster positioned within the nozzle section is provided. The method includes the steps of igniting the solid rocket booster to accelerate the aircraft to a speed at which the scramjet engine can operate, clearing the nozzle section by ejecting the solid rocket booster from the nozzle section upon exhaustion of the solid rocket booster; and injecting fuel into a flow of air passing through the scramjet engine and igniting the fuel and the air to provide thrust to the aircraft.

Semi-axisymmetric scramjet flowpath with conformal nozzle

The present invention relates to an improved variable geometry inlet for a scram jet engine having at least one combustor module. The variable geometry inlet comprises each combustor module having two sidewalls. Each of the sidewalls has a central portion with a thickness and a tapered profile forward of the central portion. The tapered profile terminates in a sharp leading edge. The variable geometry inlet further comprises each module having a lower wall and a movable cowl flap positioned forward of the lower wall. The movable cowl flap has a leading edge and the leading edges of the sidewalls intersect the leading edge of the cowl flap.

Variable geometry inlet design for scram jet engine

The present invention relates to a scram jet engine for use with a hypersonic vehicle. The scram jet engine has an upper boundary wall, a lower boundary wall, and a plurality of side walls defining an inner air flowpath. The walls are formed by a plurality of tubular heat exchanger panels. Each of the heat exchanger panels comprises a plurality of structural panels, each having a plurality of cooling passages, joined together. In the scram jet engine of the present invention, fuel may be used as a coolant and supplied to the cooling passage.

Scram jet engine design

Accordingly, the present invention provides a ducting system for use on a gas turbine engine having two fans connected to a common shaft, and includes a primary inlet duct that is integral with a plurality of primary discharge ducts, a plurality of secondary inlet ducts that are integral with a secondary discharge duct, each of the ducts includes a port, the ports of the secondary inlet ducts are located radially outward of the port of the primary inlet duct, and the ports of the primary discharge ducts are located radially outward of the port of the secondary discharge duct.

Inverter ducting for dual fan concept

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.

Flow diverter system for multiple streams for gas turbine engine

Flow diverter system for multiple streams for gas turbine engines

A fuel injector 14 passes through wall 10 to inject fuel into a hot supersonic air flow 12. Wedge seals 26, 28 seal against leakage in cooperation with injection air flow 37, being compressed within seal carrier 16. Face seals 30, 32 are compressed by cover plate 34. Inboard seals 26, 44 are preferably more porous than the outboard seals so that there is a leakage of cooling air through crevice 44.

Growing community of inventors

Hobe Sound, FL, United States of America

Daniel P Guinan