Method and arrangement for reducing the effects of a sonic boom created by an aerospace vehicle when said vehicle is flown at supersonic speed. The method includes providing the aerospace vehicle with a first spike extending from the nose thereof substantially in the direction of normal flight of the aerospace vehicle, the first spike having a second section aft of a first section that is aft of a leading end portion, the first and second sections having a second transition region therebetween and each of the sections having different cross-sectional areas, the leading end portion of the first spike tapering toward a predetermined cross-section with a first transition region between the predetermined cross-section and the first section. The first transition region is configured so as to reduce the coalescence of shock waves produced by the first spike during normal supersonic flight of the aerospace vehicle. A spike may also be included that extends from the tail of the aerospace vehicle to reduce the coalescence of shock waves produced by the spike during normal supersonic flight of the aerospace vehicle.

Supersonic Aircraft with Spike for Controlling and Reducing Sonic Boom

Embodiments of the invention relate to a supersonic nacelle design employing a bypass flow path internal to the nacelle and around the engine. By shaping the nacelle, embodiments of the invention may function to reduce sonic boom strength, cowl drag, and/or airframe interference drag. Embodiments of the invention may also function to improve total pressure recovery and/or total thrust of the primary flow path through the engine.

Low shock strength propulsion system

A supersonic inlet employs relaxed isentropic compression to improve net propulsive force by shaping the compression surface of the inlet. Relaxed isentropic compression shaping of the inlet compression surface functions to reduce cowl lip surface angles, thereby improving inlet drag characteristics and interference drag characteristics. Supersonic inlets in accordance with the invention also demonstrate reductions in peak sonic boom overpressure while maintaining performance.

Isentropic compression inlet for supersonic aircraft

A first responder support system is disclosed. The first responder support system includes a plurality of sensors deployed within a building, such that each of the plurality of sensors is configured to communicate an environmental condition, an emergency control module in communication with the plurality of sensors deployed within the building, the emergency control module is configured to receive the environmental conditions provide by the plurality of sensors, a modeling module configured to receive a building information model and to generate a building representation, and a responder support module in communication with the emergency control module and the modeling module and configured to generate an interactive rescue plan based on the building representation and the received environmental conditions.

First responder decision support system based on building information model (bim)

Active clearance control ejector systems and methods are disclosed. An example system may include an air ejector; an entrance pipe arranged to supply compressor bleed air to the air ejector; an inlet duct arranged to supply fan bypass air to the air ejector; and a supply pipe arranged to receive ejector outlet air from the air ejector and supply the ejector outlet air to an active clearance control system. The ejector outlet air may include a mixture of the compressor bleed air and the fan bypass air. The air ejector may include a venturi arranged to conduct the compressor bleed air therethrough, thereby drawing the fan bypass air into the air ejector.

System and method for active clearance control

An advanced aperture inlet (AAI) uses a three-dimensional, mixed compression inlet design derived from computational fluid dynamics (CFD) by streamline tracing a supersonic section from an axisymmetric mixed compression inlet solution. The axisymmetric design is used to obtain a CFD solution with slip wall boundaries at the inlet design point and serves as a flow field generator for the AAI. The AAI geometry is obtained by projecting a desired aperture shape onto a surface model of the external oblique shock. Streamline seeds are located on the projected aperture segments and transferred into the CFD solution space. The streamlines generated by these seeds inside the CFD solution space are then used as a wireframe to define the supersonic diffuser back to the throat location. Traditional design techniques are then used to define the subsonic diffuser from the inlet throat to the engine face.

System, method, and apparatus for designing streamline traced, mixed compression inlets for aircraft engines

A system, method, and apparatus for throat corner scoop offtake for mixed compression inlets for high speed aircraft engine applications is disclosed. The throat corner scoops are small air intakes located inside the large mixed compression inlet. They are positioned in a region otherwise prone to generate low pressure airflow. The throat corner scoops capture and remove the low pressure airflow from the bulk stream that is passed on to the engine. This location also provides inlet stability enhancement, and the airflow is used on the auxiliary systems.

System, method, and apparatus for throat corner scoop offtake for mixed compression inlets on aircraft engines

A diverterless hypersonic inlet (DHI) for a high speed, air-breathing propulsion system reduces the ingested boundary layer flow, drag, and weight, and maintains a high capture area for hypersonic applications. The design enables high vehicle fineness ratios, low-observable features, and enhances ramjet operability limits. The DHI is optimized for a particular design flight Mach number. A forebody segment generates and focuses a system of multiple upstream shock waves at desired strengths and angles to facilitate required inlet and engine airflow conditions. The forebody contour diverts boundary layer flow to the inlet sides, effectively reducing the thickness of the boundary layer that is ingested by the inlet, while maintaining the capture area required by the hypersonic propulsion system. The cowl assembly is shaped to integrate with the forebody shock system and the thinned boundary layer region.

Bradley C. Leland

Papers

System, method, and apparatus for designing streamline traced, mixed compression inlets for aircraft engines

System, method, and apparatus for throat corner scoop offtake for mixed compression inlets on aircraft engines

Method of designing a hypersonic inlet for an aircraft and system therefor

Apparatus and method for processing airflow with flowfield molded hypersonic inlet

System, method, and apparatus for diverterless hypersonic inlet for integrated turbojet and ram-scramjet applications