Aerojet Rocketdyne

Azusa, CA, United States

Aerojet Rocketdyne

Azusa, CA, United States
SEARCH FILTERS
Time filter
Source Type

Patent
Aerojet Rocketdyne | Date: 2017-05-17

There are described herein methods and systems for estimating a system parameter in a closed loop scheme using a sensor model associated with a sensor performing a measurement of the system parameter. Past and current measurements of the parameter are used to provide an initial estimate of the system parameter (606) and sensor dynamics are used to refine the estimated parameter (606).


Patent
Aerojet Rocketdyne | Date: 2017-02-08

An auxiliary power unit (10) for an aircraft includes a rotary intermittent internal combustion engine (12), a turbine (26) having an inlet in fluid communication with an outlet of the engine (12), the turbine (26) compounded with the engine (12), a compressor (20) having an inlet in fluid communication with an environment of the aircraft and an outlet in fluid communication with the aircraft, the compressor (20) rotatable independently of the turbine (26), an electric motor (64) drivingly engaged to the compressor (20), and a transfer generator (76) drivingly engaged to the engine (12), the transfer generator (76) and the electric motor (64) being electrically connected to allow power transfer therebetween. The compressor (20) or an additional compressor (21) may be in fluid communication with the inlet of the engine (12).


Patent
Aerojet Rocketdyne | Date: 2017-04-05

A turbine exhaust case (TEC) (28) of a turbofan aeroengine includes a mixer (30) for mixing exhaust gases with a bypass air stream, the TEC (28) comprising an annular hub (36) and the mixer (30) surrounding the hub (36), and a plurality of deswirling struts (58) circumferentially spaced apart with respect to a central axis (34) of the TEC (28) and located entirely within an axial length of the mixer (30). The mixer (30) defines a trailing edge (44a) having one or more upstream-most locations thereof where the mixing of the exhausted gases and the bypass air stream begins to take place. The deswirling struts (58) each extend radially across the annular exhaust gas duct (40) and interconnect the mixer (30) and the hub (36), defining a trailing edge (62) positioned upstream of and axially spaced away from the one or more upstream-most locations of the trailing edge (44a) of the mixer (30).


Patent
Aerojet Rocketdyne | Date: 2017-02-15

A gas turbine engine includes a combustor which has at least one annular wall (40, 42) defining a combustion chamber (44) therein. The annular wall (40, 42) is formed by a circumferential array of panels (60) overlapping one with another to define a plurality of radial gaps (62) between respective adjacent two panels (60). The radial gaps (62) are configured in a spiral pattern and are in fluid communication with the combustion chamber (44) and a space outside the combustor to allow air surrounding the annular wall (40, 42) to enter the combustion chamber (44) via the radial gaps (62) for film cooling.


Patent
Aerojet Rocketdyne | Date: 2017-02-08

A turboprop engine assembly (10) for an aircraft, including an internal combustion engine (12) having a liquid coolant system, an air duct (70) in fluid communication with an environment of the aircraft, a heat exchanger (66) received within the air duct (70) having coolant passages (66a) in fluid communication with the liquid coolant system and air passages (66b) in fluid communication with the air duct (70), and an exhaust duct (80) in fluid communication with an exhaust of the internal combustion engine (12). The exhaust duct (80) has an outlet (82) positioned within the air duct (70) downstream of the heat exchanger (66) and upstream of an outlet (72) of the air duct (70), the outlet (82) of the exhaust duct (80) spaced inwardly from a peripheral wall (70) of the air duct (70). In use, a flow of cooling air surrounds a flow of exhaust gases.


Patent
Aerojet Rocketdyne | Date: 2017-02-08

An auxiliary power unit (10) for an aircraft includes a rotary intermittent internal combustion engine (12) drivingly engaged to an engine shaft (16), a turbine section having an inlet in fluid communication with an outlet of the engine(s) (12), the turbine section including at least one turbine (22, 26) compounded with the engine shaft (16), and a compressor (20) having an inlet in fluid communication with an environment of the aircraft and an outlet in fluid communication with a bleed duct (72) for providing bleed air to the aircraft, the compressor (20) having a compressor rotor connected to a compressor shaft (24), the compressor shaft (24) drivingly engaged to the engine shaft (16). The driving engagement between the compressor shaft (24) and the engine shaft (16) is configurable to provide at least two alternate speed ratios between the compressor shaft (24) and the engine shaft (16).


A system and method for blade angle position feedback. The system comprises an annular member operatively connected to rotate with a propeller, a sensor fixedly mounted adjacent the annular member and configured for detecting a passage of each singularity as the annular member is rotated and axially displaced and for generating a sensor signal accordingly, the annular member and sensor configured for relative axial displacement between a first relative axial position and a second relative axial position respectively corresponding to the first and the second mode of operation, and a detection unit connected to the sensor for receiving the sensor signal therefrom, determining on the basis of the sensor signal a time interval elapsed between the passage of successive singularities, and computing from the time interval blade angle position.


A method of providing active flow control for an aircraft includes cooling a liquid coolant in a heat exchanger (28) by circulating a cooling airflow through the heat exchanger (28), and providing fluid communication between the cooling airflow and a boundary layer flow of at least one flight control surface (64; 164) of the aircraft. The cooling airflow affects the boundary layer flow of the flight control surface(s) (64; 164) to provide active flow control. A method of cooling an engine core (12) of an engine assembly (10) includes circulating a cooling fluid through the engine core (12), and cooling the cooling fluid with a cooling airflow used to provide active flow control to a flight control surface (64; 164) of the aircraft. An active flow control system (66; 166) for an aircraft is also discussed.


Patent
Aerojet Rocketdyne | Date: 2017-09-27

A propeller balancing system (216) and method that determines a stable cruise condition of an aircraft (200) during flight, and considers vibration data for propeller balancing collected while the aircraft (200) is in the stable cruise condition. Parameters used to determine stable cruise condition may be unique for each flight and/or aircraft (200), and balancing solutions may be determined based on these unique parameters.


Patent
Aerojet Rocketdyne | Date: 2017-09-20

A turbopump machine includes a housing, a shaft rotatably supported in the housing on a set of bearings, an axial pump coupled with the shaft, a circumferential discharge volute fluidly coupled with the axial pump, and a turbine coupled with the shaft. The turbine includes a blade row disposed in an axial turbine flowpath that has an axial turbine flowpath discharge that is fluidly coupled with the circumferential discharge volute. A cooling passage is disposed between the housing and the shaft about the set of bearings. The cooling passage has a cooling passage discharge that is fluidly coupled with the circumferential discharge volute. The cooling passage discharge is adjacent the axial turbine flowpath discharge. A seal isolates the cooling passage discharge from the axial turbine flowpath discharge.

Loading Aerojet Rocketdyne collaborators
Loading Aerojet Rocketdyne collaborators