Abstract: A gas turbine engine system includes a fan bypass passage in a cooling passage having an inlet that receives a bleed flow from the fan bypass passage and an outlet that discharges the bleed flow into the fan bypass passage. A nozzle having a variable cross-sectional area controls an airflow within the fan bypass passage. The bleed flow outlet is placed such that moving the nozzle to change the variable cross-sectional area controls an amount of the bleed flow through the cooling passage.

Abstract: A method, and a related material, for utilizing high performance solid rocket propellants, which are molding powders. A propellant molding powder are selected to have a design burning rate and a tailored compaction profile. A morphology of a center-port of a rocket is selected for the design burn rate and a spin-rate. The molding powder is compacted isostatically around a core through application of triaxial pressure therein forming a solid rocket propellant charge with the selected center-port shape. The solid rocket propellant charge is placed in a cartridge or a case. The cartridge is selected from various types of cartridges and specialty charges. The solid rocket propellant molding powders are highly filled with metallic fuels, and have a binder in the range of 4% to 18%, which at least partially coats the surface of the molding powder.

Abstract: A vehicle is disclosed that includes a propellant tank, an optical absorber operable to transform optical energy into thermal energy, a quantity of solid lithium within the propellant tank, a heat exchanger, and an engine. The heat exchanger is operable to transfer thermal energy from the optical absorber to the quantity of solid lithium to liquefy at least a portion of the solid lithium, and further operable to boil the liquefied portion of the solid lithium. The engine is operable to utilize lithium vapor from the boiled lithium to propel the vehicle.

Abstract: There is disclosed a field emission electric propulsion (FEEP) system including a FEEP thruster having at least one emitter and an extractor electrode, and a power supply. The power supply may provide an extractor voltage applied between the emitter and the extractor electrode. The power supply may be operable in a constant current mode in which the extractor voltage is controlled to set an ion current flowing from the emitter at a target current level.

Abstract: Dielectric barrier discharge plasma actuators are used to manipulate exhaust flow within and behind a jet engine nozzle. The dielectric barrier discharge plasma actuators may be used to direct cooling airflow near the surface of the nozzle to reduce heating of the nozzle, create thrust vectoring, and reduce noise associated with the exhaust flow exiting the nozzle.

Abstract: A nano-particle field extraction system comprising a grid having a plurality of electrodes each defining an electrical field, wherein the grid has a plurality of vias extending therethrough. The system further comprises a reservoir having a generally dry mixture disposed therein, a plurality of particles suspended in the generally dry mixture, a biasing member applying a biasing force to the generally dry mixture in the reservoir, and a sieve electrode system in electrical communication with the grid. The sieve electrode system has a plurality of through-holes extending from the reservoir to the grid, such that the sieve electrode system cooperates with the biasing member to extract at least one particle from the generally dry mixture and into the grid whereby the electrical fields charge and accelerate the particle in the vias.

Abstract: An electrospray source useful for a variety of applications and including an emitter with a porous media flow distributor having a surface forming multiple Taylor cones. A casing about the porous media flow distributor controls the direction of a working fluid through the porous media. An extractor is at a potential different than the emitter for forming the Taylor cones. A guard electrode is disposed between the emitter and the extractor and is at or above the potential of the emitter for shaping the electric field formed between the emitter and the extractor.

Abstract: A rocket engine fluid-flow system includes a pump fluidly interconnecting a fluid source to a combustion chamber. A nozzle is in fluid communication with the combustion chamber and includes coolant tubes fluidly arranged between the pump and the combustion chamber. An orifice has a throat and is fluidly arranged between the pump and the coolant tubes. The orifice has entrance and exit ramps arranged on either side of the throat. The exit ramp has an exit ramp surface with a divergent angle that is less than a right angle. The entrance ramp provides a smooth approach to the orifice throat. In one example, the exit ramp includes an exit ramp surface having a divergent angle of 20-60°. The exit ramp radius is less than twice the throat radius in one example.

Abstract: An engine contains at least one pulse detonation combustor having a combustion chamber and an exit nozzle coupled to and downstream of the combustion chamber. During operation of the at least one pulse detonation combustor a detonation occurs within the combustion chamber and at least one of a fuel fill fraction and purge fraction of the at least one pulse detonation combustor are utilized to offset a temperature peak of said detonation from a pressure peak of said detonation. The fuel fill fraction is defined as 1?purge fraction, and the purge fraction is the ratio of the purge time of the at least one pulse detonation combustor to a sum of the purge time of the at least one pulse detonation combustor and a fuel fill time of the at least one pulse detonation combustor.

Abstract: A gas turbine engine with a two-spool fan and a method of operating the gas turbine engine includes modulating a pitch of a variable pitch first stage fan section of a low pressure spool, the variable pitch first stage fan section upstream of a second stage fan section of an intermediate pressure spool to maintain a generally constant engine inlet flow while modulating engine thrust.

Abstract: A gas turbine engine and a method of operating the gas turbine engine according to an exemplary aspect of the present disclosure includes modulating a variable high pressure turbine inlet guide vane of a high pressure spool to performance match a first stage fan section of a low pressure spool and an intermediate stage fan section of an intermediate spool to maintain a generally constant engine inlet flow while varying engine thrust.

Abstract: An expander cycle rocket engine includes secondary turbopump to further pressurize a gaseous fuel discharged from a primary turbine prior to entering the combustion chamber. The secondary turbopump is driven by fuel bled off from the primary fuel pump. A gaseous fuel that is heated from passing around the nozzle that is passed through the primary turbine to drive the primary fuel and oxidizer pumps is then passed through the secondary turbine to drive the secondary fuel compressor. With the secondary turbopump used in the Johnson-Sexton cycle engine, a thrust produced by the expander cycle rocket engine is greater than those obtained by prior art expander cycle rocket engines due to the square-cube rule.

Abstract: An injector includes an injector body that has at least an injection side surface. A plurality of passages extends within the injector body for injecting combustion fluids at the injection side surface. The plurality of passages include a first set of passages that has a first arrangement defining a first impingement point in space with regard to the injection side surface and a second, different set of passages that has a second arrangement defining a second impingement point in space with regard to the injection side surface.

Abstract: A method operable to improve pressure recovery and/or distortion within engine inlet is disclosed. A first fluid flow is provided to primary jet vortex generator(s) operable to inject fluid at a first injection rate into a boundary layer of a primary fluid flow within the inlet. A secondary fluid flow is injected by secondary jet vortex generator(s) at a second injection rate into the boundary layer of the primary fluid flow. The fluid injected at the first injection rate and second injection rate is operable to induce secondary flow structures within the boundary layer. These secondary close structures are then operable to improve or manipulate the pressure recovery of the inlet. At specific engine conditions, this method may redistribute the ratio of the first injection rate and second injection rate in order to improve pressure recovery and/or distortion of the inlet when the particular engine conditions.

Abstract: A thrust reverser includes a slider movable along an actuator shaft. An inner linkage is mounted to the slider and the inner thrust reverser door and an outer linkage is mounted to the slider and the outer thrust reverser door.

Abstract: A method of controlling mixing of a flow exiting a downstream end of a primary nozzle associated with a jet engine. The method may involve coupling a shape memory alloy (SMA) element to a mixing structure disposed at the downstream edge of the primary nozzle. An electrical signal may be applied to the SMA element to heat the SMA element and induce a phase change in the SMA element. The phase change may cause an axial length of the SMA element to constrict, to cause movement of the mixing structure into a path of the flow exiting the primary nozzle.

Abstract: Disclosed are embodiments for producing thrust, and in particular thrust for the propulsion of a flight vehicle. The embodiments incorporate the on-board laser heating of a propellant to a plasma state for the production of thrust, and of energy being supplied to the on-board laser by remote power sources such as ground based, sea based, space based, or airborne pump lasers.

Abstract: The invention aims at providing a sufficient air exchange in the engine area without raising the static pressure in the outlet area of the air exchange duct. In order to do this, the invention provides for an outlet section capable of concentrating the outlet of the exchange airflow towards zones with smaller static pressure. An turbojet engine outlet according to the invention provides an engine cowl, which comprises an external wall of a circumferential duct in which circulates an exchange airflow of the turbines, and an airflow mixer. At the outlet, the duct delivers the exchange airflow towards the mixer for the primary and secondary airflows of the turbojet engine. The mixer presents hot and cold lobes. In particular, the engine outlet also comprises guiding means of the exchange airflow towards the cold lobes.

Abstract: Closed-cycle rocket engine assemblies including a combustor assembly, a combustor jacket, a turbine, a first pump and a mixing chamber are disclosed. The combustor jacket facilitates the transfer of heat from the combustor assembly into a fluid and the turbine is driven by a heated fluid from the combustor jacket. The mixing chamber may include a first inlet to receive a fluid from the turbine, a second inlet to receive a fluid from a first reactant reservoir, and an outlet to deliver a fluid to the first pump. Additionally, the first pump may be coupled to and powered by the turbine and the first pump may be configured to deliver at least a portion of the fluid from the mixing chamber into the combustion chamber of the combustor assembly. Related methods of operating such rocket engine assemblies are also disclosed.

Abstract: A method for starting up an aircraft turbine engine using an electric starter is provided. The aircraft turbine engine is provided with at least one hydraulic fluid circuit for lubricating the aircraft turbine engine, the aircraft turbine engine lubricating circuit is provided with a temperature-regulating section that is able to remove the heat generated by the aircraft turbine engine when the aircraft turbine engine is running, the electric starter is provided with at least one hydraulic fluid circuit for lubricating parts of the electric starter, and the lubricating circuit of the electric starter is provided with the temperature-regulating section that is able to remove the heat emitted by the electric starter when the lubricating circuit of the electric starter is operating.

Abstract: A method for reducing the vibration levels likely to occur, in a turbine engine including a first and a second bladed disk forming a propfan of contrarotating disks, when the two disks are traversed by a gaseous fluid, because of the turbulence of aerodynamic origin generated by the second bladed disk on the first bladed disk is disclosed. The method includes defining an initial configuration of the blades, calculating the synchronous forced response on the first bladed disk as a function of the harmonic excitation force generated by the second bladed disk expressed as a linear function of the generalized aerodynamic force for the mode in question; for stacked sections of one of the two disks, determining a tangential geometric offset value ? of the individual aerodynamic profile to reduce the term corresponding to the generalized aerodynamic force; and applying a new configuration to the blades.

Abstract: (A1) A turbofan engine (10) is provided that includes a spool (14). The spool (14) supports a turbine (18) and is housed within a core nacelle (12). A fan (20) is coupled to the spool (14) and includes a target operability line. The target operability line provides desired fuel consumption, engine performance, and/or fan operability margin. A fan nacelle (34) surrounds the fan (20) and core nacelle (12) to provide a bypass flow path (39) having a nozzle exit area (40). A controller (50) is programmed to command a flow control device (41) for changing the nozzle exit area (40). The change in nozzle exit area (40) achieves the target operability line in response to an engine operating condition that is a function of airspeed and throttle position. A change in the nozzle exit area (40) is used to move the operating line toward a fan stall or flutter boundary by manipulating the fan pressure ratio.

Abstract: The invention relates to a nacelle for a variable area nozzle propulsion unit comprising a nacelle, hosting a turbofan jet engine, of dual-flow type. The propulsion unit includes: a fixed part, carried by the inboard half-nacelle defined by a vertical plane of symmetry of the nacelle; and a movable part carried by the outboard half-nacelle. The moving part containing or releasing part of the secondary flow, depending on its open or closed position. The moving part being able to move to a discrete number of positions including a closed position, an open position, and one or more intermediate positions so as to provide variable area nozzle configurations for the turbofan engine.

Abstract: A pulsed detonation engine may include a detonation tube for receiving fuel and an oxidizer to be detonated therein, one or more fuel-oxidizer injectors for injecting the fuel and oxidizer into the detonation tube, one or more purge air injectors for injecting purge air into the detonation tube for purging the detonation tube, and an ignition for igniting the fuel and oxidizer in the detonation tube so as to initiate detonation thereof. The detonation tube has an upstream end, a downstream end, and an axially extended portion extending from the upstream end to the downstream end and having a perimeter. The fuel-oxidizer injectors and purge air injectors may be disposed at least along the axially extended portion. The ignition may include a plurality of igniters disposed at or near the perimeter of the axially extended portion, spaced about the perimeter, at or near the upstream end of the detonation tube.

Abstract: A method of using one or more microjets to create and/or control oblique shock waves. The introduction of microjet flow into a supersonic stream creates an oblique shock wave. This wave can be strengthened—by increasing microjet flow rate or the use of many microjets in an array—in order to form an oblique shock. Such an oblique shock can be used to decelerate flow in a jet aircraft engine inlet in a controlled fashion, thus increasing pressure recovery and engine efficiency while reducing flow instability. Adjusting the pressure ratio across the microjet actually alters the angle of the oblique shock. Thus, the use of microjets allows decelerating shock waves in an inlet engine to be properly positioned and controlled. Microjet arrays can also be used to ameliorate shock waves created by external aircraft surfaces, such as sensor pods and weapons.

Abstract: A method of cooling a rocket engine that includes removal of a gaseous propellant from a tank containing liquid propellant, thereby cooling the propellant in the tank, cooling a rocket engine with a coolant, thereby heating the coolant, and transferring heat from the heated coolant into the propellant in the tank, thereby heating the propellant in the tank and maintaining the pressure in the tank. In certain embodiments, the coolant is injected into the rocket engine after transferring heat from the rocket engine to the propellant in the tank.

Abstract: A gas turbine engine system is disclosed which includes a core passage and a bypass passage which can be configured as a fan bypass duct or a third stream bypass duct. The core passage and bypass passage are routed to flow through a nozzle before exiting overboard an aircraft. The nozzle includes moveable members capable of changing a configuration of the nozzle. In one form the moveable members are capable of changing throat area for portions of the nozzle that receive working fluid from the core passage and the bypass passage. The bypass passage can include a branch. In one form the branch can include a heat exchanger. The bypass passage can also provide cooling to one or more portions of the nozzle, such as cooling to a deck of the nozzle.

Abstract: A constant volume combustion chamber, combustor, and method for constant volume combustion involve combusting a fuel in a chamber sealed by a pintle having a conical portion fitted into a conical nozzle throat and pulling the pintle away from the nozzle throat to allow combustion products to exhaust through a nozzle outlet. The shapes and surfaces of the pintle and nozzle throat provide for sealing the chamber at high pressures while resisting surface wear. Operational parameters for the combustor may be computer controlled in response to measured pressures and temperatures in the combustor.

Abstract: 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.

Abstract: 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.

Abstract: An engine and a method of operating the engine are provided. The engine includes a gas turbine and fan that rotate together to provide an exhaust gas flow stream, which flows over a free turbine that is connected to a power take-off. The free turbine can extract energy from the exhaust gas flow stream and transfer it as shaft power to the power take-off and the amount of energy extracted by the free turbine is controlled by varying the pitch of the free turbine's blades and/or by varying the pitch or stator vanes of a stator upstream of the free turbine. The control over the amount of energy extracted by the free turbine allows the engine to be used to provide thrust from the gas turbine and fan or to provide shaft power at the power take-off, or a combination of thrust and shaft power.

Abstract: An apparatus for reducing heating effects of an exhaust plume of a jet engine on an impinged surface includes fluid injectors disposed adjacent and aimed into an exhaust plume zone that's to be occupied by an exhaust plume when the engine is running. A flow generator transmits fluid flow into such an exhaust plume through the injectors. Each injector emits fluid in at least two divergent directions to increase the cross-sectional area of the exhaust plume by forming fluidic lobes in the exhaust plume.

Abstract: (A1) A turbofan engine control system is provided for managing a fan operating line. The engine (10) includes a spool having a turbine housed in a core nacelle (12). A turbofan (20) is coupled to the spool (14). A fan nacelle (34) surrounds the turbofan and core nacelle and provides a bypass flow path having a nozzle exit area (40). A controller (50) is programmed to effectively change the nozzle exit area in response to an undesired turbofan stability margin which may result in a stall or flutter condition. In one example, the physical nozzle exit area is increased at the undesired stability condition in which the airflow into the engine creates a destabilizing pressure gradient at the inlet side of the turbofan. A turbofan pressure ratio, turbofan pressure gradient, low spool speed and throttle position are monitored to determine the undesired turbofan stability margin.

Abstract: Propellants flow through specialized mechanical hardware that is designed for effective and safe ignition and sustained combustion of the propellants. By integrating a micro-fluidic porous media element between a propellant feed source and the combustion chamber, an effective and reliable propellant injector head may be implemented that is capable of withstanding transient combustion and detonation waves that commonly occur during an ignition event. The micro-fluidic porous media element is of specified porosity or porosity gradient selected to be appropriate for a given propellant. Additionally the propellant injector head design integrates a spark ignition mechanism that withstands extremely hot running conditions without noticeable spark mechanism degradation.

Abstract: An aircraft propulsion arrangement including a gas turbine aircraft engine having a compressor, an oil system configured to route engine oil through a heat exchanger mounted to an external surface of an aircraft component so as to define part of an aerodynamic surface to a flow of ambient air, and a duct arrangement fluidly connecting the compressor to the heat exchanger. A method of operating the gas turbine aircraft engine, the method involving the steps of: (a) flowing engine oil through the heat exchanger and thus into heat-exchange relationship with the ambient air; and (b) directing a bleed flow of compressor gas drawn from the compressor along the duct arrangement and into heat-exchange relationship with the engine oil in the heat exchanger, wherein the directing step (b) is performed selectively.

Abstract: 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.

Abstract: A gas turbine engine system includes a fan bypass passage (27), a core nacelle (28) having an inner fixed structure (40) within the fan bypass passage, a passage (42) extending through the inner fixed structure, and a duct nozzle (48). The passage includes an inlet (44) for receiving a fan airflow (F2) from the fan bypass passage and an outlet (46) for discharging the fan airflow. The duct nozzle includes a variable cross-sectional exit area (50) for controlling the fan airflow within the passage and is selectively moveable to influence the variable cross-sectional exit area.

Abstract: The invention relates to a method for controlling at least one actuator for actuating the cowlings of a thrust inverter in a turbojet engine, the actuator being driven by an electric motor including a relative position sensor providing information on the evolution of the movement thereof, wherein the motor is controlled based on the instantaneous position of the cowling in at least one portion of the movement thereof between an open position and a closed position, the instantaneous position of the cowling being determined from at least one reference position absolute data and relative position data relative to said reference position provided by the relative position sensor of the motor, wherein in case the actuation is resumed after an interruption, a new determination of the reference position is initiated.

Abstract: A propulsion system for a combined aircraft/spacecraft including: a jet engine having a longitudinal axis defined in a direction from a front end at an air intake to a rear end at a jet exhaust: a front end at an air intake to a rear end at the jet exhaust; a rocket engine having a longitudinal axis defined normal to a rear end at a rocket exhaust; a common engine housing having an elongated shape and wherein the jet engine and the rocket engine are configured substantially coaxially and with respective rear ends facing the same direction when the rocket engine is operated.

Abstract: A propulsion system may be operated by determining pressure in a cryogenic liquid tank storing a fluid and cooling the cryogenic liquid tank in response to determining that the pressure is greater than a predetermined value. The cryogenic liquid tank may be pressurized by admitting a gaseous form of the fluid into the cryogenic liquid tank in response to determining that the pressure in the cryogenic liquid tank is less than a predetermined value.

Abstract: A vehicle propulsion system having a manifold in fluid communication with a cross flow fan which adjusts to control an amount lift generated by a plurality of airfoils providing lift from air from the cross flow fan.

Abstract: A turbofan engine (10) includes a fan variable area nozzle (FVAN) (50) which effectively changes the physical area and geometry within a fan bypass flow path (40) to manipulate the pressure ratio of the bypass flow. The FVAN generally includes a multitude of aerodynamically shaped inserts (52) circumferentially located about the core nacelle (12). The FVAN at a fully stowed position (takeoff/landing) takes up a minimum of area within the fan bypass flow path to effectively maximize the fan nozzle exit area (44) while in the fully deployed position (cruise) takes up a maximum of area within the bypass flow path to effectively minimize the fan nozzle exit area. By separately adjusting each of the multiple of inserts relative the other inserts the FVAN provides an asymmetrical fan nozzle exit area to selectively vector the fan bypass flow.

Abstract: A system and method for assessing the risk of conjunction of a rocket body with orbiting and non-orbiting platforms. Two-body orbital dynamics are used to initially determine the kinematic access for a ballistic vehicle. The access may be represented in two ways: as a volume relative to its launcher and also as a geographical footprint relative to a target position that encompasses all possible launcher locations.

Abstract: A rocket multi-nozzle grid assembly includes an insulator grid, a plurality of refractory nozzle fittings and a grid support plate. The grid support plate braces the plurality of refractory nozzle fittings and the insulator grid. The insulator grid includes insulator orifices, each of the nozzle fittings includes a fitting orifice and the grid support plate includes a plurality of plate orifices. The rocket multi-nozzle grid assembly includes a plurality nozzles. Each nozzle includes a plate orifice aligned with one insulator orifice and one fitting orifice. Exhaust gases are directed through the plurality of nozzles. The multi-nozzle grid assembly substantially maintains its surface area throughout operation of a rocket motor. In one example, the surface area is maintained through ablation of the insulator grid. In another example, the surface area is maintained through cooling of the plurality of nozzle fittings with the ablated fragments passing through the nozzles.

Abstract: Systems and methods involving multiple torque paths of gas turbine engines are provided. In this regard, a representative spool assembly for a gas turbine engine, which incorporates a compressor, a turbine and a gear assembly, includes: a shaft operative to be driven by the turbine; a first spool segment operative to couple the shaft to the compressor; and a second spool segment operative to couple the shaft to the gear assembly. The first spool segment and the second spool segment are not coupled to each other.

Abstract: A system and methods are provided for combining systems of an upper stage space launch vehicle for enhancing the operation of the space vehicle. Hydrogen and oxygen already on board as propellant for the upper stage rockets is also used for other upper stage functions to include propellant tank pressurization, attitude control, vehicle settling, and electrical requirements. Specifically, gases from the propellant tanks, instead of being dumped overboard, are used as fuel and oxidizer to power an internal combustion engine that produces mechanical power for driving other elements including a starter/generator for generation of electrical current, mechanical power for fluid pumps, and other uses. The exhaust gas from the internal combustion engine is also used directly in one or more vehicle settling thrusters. Accumulators which store the waste ullage gases are pressurized and provide pressurization control for the propellant tanks.

Abstract: A nozzle device defines a passageway including an outlet to discharge working fluid to produce thrust. This device includes a vectoring mechanism having three or more vanes pivotally mounted across the passageway and a linkage pivotally coupling the vanes together. This linkage includes a first arm fixed to a first one of the vanes to pivot therewith about a first pivot axis, a second arm and a third arm fixed to a second one of the vanes to pivot therewith about a second pivot axis, and a fourth arm fixed to a third one of the vanes to pivot therewith about a third pivot axis. A first connecting link pivotally couples the first arm and the second arm together, and a second connecting link pivotally couples the third arm and the fourth arm together. The relative angular positioning of the arms with respect to the corresponding pivot axes and/or the arm links can be varied to define different vectoring schedules with the mechanism linkage.

Abstract: In one embodiment, a pulse detonation engine includes a resonator configured to fluidly couple to an air flow path upstream of a pulse detonation tube. The pulse detonation engine also includes a controller configured to receive signals indicative of an operating frequency of an air valve disposed at an upstream end of the pulse detonation tube, and to adjust a geometric configuration of the resonator in response to the signals.

Abstract: A method for producing the electrical power required for equipment on board an airplane is disclosed. Auxiliary power is taken off by a shaft driven by the high pressure turbine and, when idling, the efficiency of the low pressure turbine is degraded so as to enable the high pressure turbine to operate at a speed that is sufficient for delivering the required auxiliary power.

Abstract: There is disclosed a vehicle and methods for maneuvering the vehicle. The vehicle may include a plurality of multiple-impulse rocket motors, each of which comprises a plurality of independently ignitable solid fuel propellant charges, and a processor that generates at least one command to ignite at least one solid fuel propellant charge of at least one of the plurality of multiple-impulse rocket motors.