Showing papers on "Rocket published in 1994"

Hot-gas-side and coolant-side heat transfer in liquid rocket engine combustors

Abstract: The objectives of this article are to develop a multidisciplinary, computational methodology to predict the hot-gas-side and coolant-side heat transfer in film cooling assisted, regeneratively cooled liquid rocket engine combustors, and to use it in parametric studies to recommend optimized design of the coolant channels for a developmental combustor. An integrated numerical model which incorporates computational fluid dynamics (CFD) for the hot-gas thermal environment, and thermal analysis for the liner and coolant channels, was developed. This integrated CFD/thermal model was validated by comparing predicted heat fluxes with those of hot-firing test and industrial design methods for a 40-k calorimeter thrust chamber and the Space Shuttle Main Engine main combustion chamber. Parametric studies were performed for the advanced main combustion chamber to find a strategy for a proposed coolant channel design.

Solid rocket exhaust in the stratosphere - Plume diffusion and chemical reactions

Abstract: A model has been developed to examine, on a local scale, the reactions of rocket exhaust from solid rocket motors with stratospheric ozone. The effects were examined at two different altitudes. Results of the modeling study indicate that afterburning chemistry of reactive exhaust products can cause local but transient (on the order of several minutes) loss of ozone. The modeling study included potential heterogeneous reactions at aluminum oxide surfaces. Results indicate that these potential heterogeneous reactions do not have a major impact on the local plume chemistry. Homogeneous reactions appear to be of more consequence during the early dispersion of the plume. It has also been found that the rate of plume dispersion has a very significant effect on local ozone loss.

Damage-Mitigating Control of a Reusable Rocket Engine

Abstract: This article presents a control concept for damage prediction and damage mitigation in reusable rocket engines for enhancement of structural durability The key idea here is to achieve high performance without overstraining the mechanical structures so that 1) the functional lives of critical components are increased, resulting in enhanced safety, operational reliability, and availability; and 2) the plant (ie, the rocket engine) can be inexpensively maintained, and safely and efficiently driven under different operating conditions To this effect, dynamics of fatigue damage have been modeled in the continuous-time setting instead of the conventional cycle-based approach, and an optimal control policy is formulated by constraining the accumulated damage and its time derivative Efficacy of the proposed damage mitigation concept is evaluated for life extension of the turbine blades of a bipropellant rocket engine via simulation experiments The simulation results demonstrate the potential of increasing the structural durability of reusable rocket engines with no significant loss of performance

Electron time dispersion

Abstract: Observations of electrons during a sounding rocket flight from Sondrestromfjord, Greenland, in 1985 showed numerous instances of electron time-of-flight dispersion, that is, a coherent structure in which fast electrons arrived before slow electrons. The instrumentation measured many pitch angles simultaneously with high time resolution, providing a detailed view of the temporal evolution of the electron signature at rocket altitudes which was indicative of an impulsive injection of electrons over a range of energies. A model utilizing an extended, quasi-static (static on a timescale of 1 s), parallel electric field was developed to investigate a possible mechanism for the injection of these electrons. The model is found to fit the data well only when a background is added which cannot be self-consistently generated by the model, although other parameters used in the model are in general agreement with previous reports. This leads to the conclusion that electron dispersion is more probably generated by a wave acceleration mechanism.

Launch vehicle system

Abstract: A launch vehicle system includes at least one primary solid fuel motor, a shroud for housing a payload and an attitude control system. The attitude control system is connected intermediate the primary solid fuel motor and the shroud. It has a circular outer wall and liquid fuel motors for exerting thrust along a longitudinal axis of the rocket system, and for controlling pitch, roll and yaw of the shroud. The attitude control system includes a support structure defining three pairs of fuel tank supporting zones, each zone having laterally extending supporting brackets connected to the outer wall. The attitude control system also includes connecting tubing for connecting outlets of sets of fuel tanks consisting of one, two and three pairs of fuel tanks to the liquid fuel motors.

Pressurized air/water rocket and launcher

Abstract: The present invention involves a liquid jet propelled rocket and rocket launcher. The launcher has a housing which includes a vessel for holding pressurized air therein, an inlet to the vessel and an outlet from the vessel. Also, the housing has a jet tube receiver extending from the outlet and adapted to connect with a jet tube of a rocket assembly. A pump is connected to the vessel inlet of the housing, the pump is connected for and capable of pumping air into the vessel at a pressure sufficient to launch the rocket assembly. A one way valve is connected to the pump and permits the flow of air only from the pump to the vessel. There is a rocket assembly latch mechanism located on the housing with means for releasing the latch. There are also, a rocket assembly which includes a liquid reservoir for receiving liquid and subsequently receiving air under pressure from the pump, a jet tube extending from the liquid reservoir and adapted to sealably and releasably connect to said jet tube receiver of the housing. The rocket assembly with the jet tube is releasably attachable to the housing with the jet tube coupled in fluid communication with the jet tube receiver by the latch mechanism, wherein the liquid may be stored within the rocket reservoir, and air may be pumped into the rocket reservoir and into the launcher vessel by means of actuating the pump.

System Sensitivity Analysis Applied to the Conceptual Design of a Dual-Fuel Rocket SSTO

Abstract: This paper reports the results of initial efforts to apply the System Sensitivity Analysis (SSA) optimization method to the conceptual design of a single-stage-to-orbit (SSTO) launch vehicle. SSA is an efficient, calculus-based MDO technique for generating sensitivity derivatives in a highly multidisciplinary design environment. The method has been successfully applied to conceptual aircraft design and has been proven to have advantages over traditional direct optimization methods. The method is applied to the optimization of an advanced, piloted SSTO design similar to vehicles currently being analyzed by NASA as possible replacements for the Space Shuttle. Powered by a derivative of the Russian RD-701 rocket engine, the vehicle employs a combination of hydrocarbon, hydrogen, and oxygen propellants. Three primary disciplines are included in the design - propulsion, performance, and weights & sizing. A complete, converged vehicle analysis depends on the use of three standalone conceptual analysis computer codes. Efforts to minimize vehicle dry (empty) weight are reported in this paper. The problem consists of six system-level design variables and one system-level constraint. Using SSA in a 'manual' fashion to generate gradient information, six system-level iterations were performed from each of two different starting points. The results showed a good pattern of convergence for both starting points. A discussion of the advantages and disadvantages of the method, possible areas of improvement, and future work is included.

The rationale/benefits of nuclear thermal rocket propulsion for NASA's lunar space transportation system

Abstract: The solid core nuclear thermal rocket (NTR) represents the next major evolutionary step in propulsion technology With its attractive operating characteristics, which include high specific impulse (approximately 850-1000 s) and engine thrust-to-weight (approximately 4-20), the NTR can form the basis for an efficient lunar space transportation system (LTS) capable of supporting both piloted and cargo missions Studies conducted at the NASA Lewis Research Center indicate that an NTR-based LTS could transport a fully-fueled, cargo-laden, lunar excursion vehicle to the Moon, and return it to low Earth orbit (LEO) after mission completion, for less initial mass in LEO than an aerobraked chemical system of the type studied by NASA during its '90-Day Study' The all-propulsive NTR-powered LTS would also be 'fully reusable' and would have a 'return payload' mass fraction of approximately 23 percent--twice that of the 'partially reusable' aerobraked chemical system Two NTR technology options are examined--one derived from the graphite-moderated reactor concept developed by NASA and the AEC under the Rover/NERVA (Nuclear Engine for Rocket Vehicle Application) programs, and a second concept, the Particle Bed Reactor (PBR) The paper also summarizes NASA's lunar outpost scenario, compares relative performance provided by different LTS concepts, and discusses important operational issues (eg, reusability, engine 'end-of life' disposal, etc) associated with using this important propulsion technology

Axial thrust behavior in LOX-pump of rocket engine

Abstract: The LOX pump of the first stage of the H-II rocket, the next generation of large launch vehicle in Japan, has shown fairly good axial thrust performance. However, the behavior of the axial thrust is not well known because of the complicated mechanism of the thrust-balancing device. In order to elucidate the flow characteristics of the complicated thrust balancing device and to improve it, the internal flow in the device was fully analyzed by developing a method of boundary value determination at each element composing the device. The analysis developed here was confirmed to give satisfactory results by comparison with actual measurements. Using the present analysis, the axial thrust behavior of the LOX pump was revealed for various combinations of balancing piston, balancing holes, swirl breaker, etc.

Techniques for developing approximate optimal advanced launch system guidance

Abstract: In this paper an extension to our previous technique used to develop a real-time guidance scheme for the Advanced Launch System is presented. Our approach is to construct an optimal guidance law based upon an asymptotic expansion associated with small physical parameters, ?. The problem is still to maximize the payload into orbit subject to the equations of motion of a rocket over a nonrotating spherical Earth. The trajectory of a rocket modeled as a point mass is considered with the flight restricted to an equatorial plane while reaching an orbital altitude at orbital injection speeds. The dynamics of this problem can be separated into primary effects due to thrust and gravitational forces, and perturbation effects which include the aerodynamic forces and the remaining inertial forces. An analytic solution to the reduced-order problem represented by the primary dynamics is possible. The Hamilton-Jacobi-Bellman or dynamic programming equation is expanded in an asymptotic series where the zeroth-order term (? = 0) can be obtained in closed form. The neglected perturbation terms are included in the higher-order terms of the expansion. These higher-order terms are determined from the solution of first-order linear partial differential equations requiring only integrations which are quadratures. These quadratures can be performed rapidly with the emerging computer capability so that real-time approximate optimization can be used to construct the launch guidance law. Here ? is chosen as the ratio of the atmospheric scale height to the radius of the Earth. It is important that the perturbation effects remain small.

Miniscale ballistic motor testing method for rocket propellants

Abstract: A method for using a miniscale ballistic test motor for determining burn es over an operating pressure range allows the testing of a small propellant sample. The small propellant sample allows performance of an abbreviated procedure for each test of the propellant sample involving loading of a small scale test motor with the sample, conditioning the test motor and sample therein, firing the propellant sample and recording data.

Hybrid solar rocket utilizing thermal storage for propulsion and electrical power

Abstract: A solar rocket for propelling and powering the electronics of a spacecraft includes a black body cavity of thermal storage material. Propulsion tubes and a connected nozzle axially extend through the black body cavity for burning propellant in order to provide thrust for the rocket. An insulation sleeve is removably located at the outer periphery of the cavity of thermal storage material. Energy conversion diodes surround the insulation sleeve. The thermal storage material, the insulation sleeve and the energy conversion diodes form a receiver. The receiver has a pair of spaced apart holes therethrough which lead into an interior space of the black body cavity in order to receive sunlight. A mirror assembly is located near each hole of the receiver for harnessing sunlight and providing radiant energy through the holes into the internal space of the black body cavity.

Rocket deployment system for parachutes and the like

Abstract: The invention is a compressed gas rocket apparatus for deployment of emergency parachutes, rescue lines and similar payloads. A pressurized vessel is equipped with a stopping or sealing mechanism which, when removed or punctured, causes the pressure vessel to be launched. A drag line connected to the vessel pulls the payload away from the launch point, in the direction of the pressure vessel travel.

Method of simultaneously forming rocket thrust chamber cooling tubes

Abstract: A plurality of tubes are placed side-by-side on the forming surface of a die, another die is brought into contact with the tubes, thus sandwiching the tubes therebetween. The two dies are then placed into a press that forces the dies together, simultaneously deforming the tubes into intimate contact with each other, thereby producing an exacting fit between adjacent tubes.

HVOF gas flow field characteristics

Abstract: The effects of combustion chamber pressure and fuel/oxygen mixture ratio on the characteristics of a high pressure, supersonic HVOF gun are examined experimentally and theoretically. The measured temperature, velocity and entrained air fraction are obtained from an enthalpy probe/mass spectrometer system. Predictions of combustion chamber flame temperature and composition are calculated with an equilibrium combustion model. Nozzle and barrel exit conditions are calculated using a one-dimensional rocket performance model. The calculations are bounded by the assumption of frozen and equilibrium compositions. Comparisons between measurements and the predictions indicate that the flow field is far from chemical equilibrium. The aerodynamic force available for accelerating a particle is primarily controlled by the chamber pressure while the composition and temperature of the gas surrounding the particles is controlled by the mixture ratio.

Technology and staging effects on two-stage-to-orbit systems

Abstract: Horizontal takeoff and landing two-stage systems with an airbreathing first stage and rocket second stage are evaluated for staging Mach numbers that range from 5 to 14. All systems are evaluated with advanced technologies being developed in the NASP Program and sized to the same mission requirements. With these advanced technologies, the two-stage systems are heavier than the single stage. The weights of the two-stage systems are closely related to staging. Using a rocket on the first stage to accelerate from the turboramjet limit of Mach 6 to Mach 10 signiificantly decreases dry weight as compared to the Mach 6-staged system. The optimum dry weight staging Mach number for the scramjet two-stage system is Mach 12. At a 40 percent weight growth (current technology level), the scramjet two-stage systems are half the weight and less sensitive to weight changes than the single stage, but still require substantial technology development in the areas of inlets, nozzles, ramjet propulsion, active cooling, and high-temperature structures.

Method of making cooled rocket combustion chamber construction

Abstract: A method of producing a rocket combustion chamber/nozzle assembly in which a liner blank is deformed against the forming surface of a die to form a finished cooling liner, a structural jacket is electro-formed about the finished cooling liner, and fluid manifolds are secured to the finished cooling liner and structural jacket assembly

Thruster firing effects in the shuttle environment: 1. Neutral gas composition

Abstract: The changes in the composition of the neutral gas in the space shuttle payload bay caused by the shuttle's vernier reaction control system rocket engines were measured with a quadrupole mass spectrometer aboard STS-4. The average hydrogen level during thruster firings was approximately 25 times greater than nitric oxide, NO, which was previously found to be one of the largest components of the gas in the payload bay during firings. The attitude of the orbiter also influenced the effects of the vernier engines as measured in the payload bay. When the mass spectrometer was pointed into the velocity vector, thrusters whose exhaust plumes were directed approximately upstream into the flow of ambient neutral species caused decreases in atomic oxygen signals. NO was enhanced in the payload bay during thruster firings primarily when the shuttle was in a bay-to-ram attitude. The return flux from the left-side, downward pointing vernier thruster was detected by the mass spectrometer only when the shuttle was in a belly-to-ram attitude. All of these results are consistent with earlier observations that the kinematics of gas phase and surface scattering are important in determining the effects of the thruster firings on the payload bay environment.

Fundamental phenomena on fuel decomposition and boundary layer combustion processes with applications to hybrid rocket motors

Abstract: This final report summarizes the major findings on the subject of 'Fundamental Phenomena on Fuel Decomposition and Boundary-Layer Combustion Processes with Applications to Hybrid Rocket Motors', performed from 1 April 1994 to 30 June 1996 Both experimental results from Task 1 and theoretical/numerical results from Task 2 are reported here in two parts Part 1 covers the experimental work performed and describes the test facility setup, data reduction techniques employed, and results of the test firings, including effects of operating conditions and fuel additives on solid fuel regression rate and thermal profiles of the condensed phase Part 2 concerns the theoretical/numerical work It covers physical modeling of the combustion processes including gas/surface coupling, and radiation effect on regression rate The numerical solution of the flowfield structure and condensed phase regression behavior are presented Experimental data from the test firings were used for numerical model validation

Shuttle launch system for model rocket

Abstract: A shuttle launch system for a model rocket which includes a model rocket engine having a propellant charge for propelling the model rocket in a first direction and an ejection charge for ejecting matter in a second direction generally opposite the first direction consists of a booster rocket having a longitudinally extended tube with a peripheral sidewall, an inner chamber and top and bottom ends and an engine-receiving chamber adjacent the bottom end of the tube. A shuttle glider support structure is mounted on the booster rocket for releasably supporting a shuttle glider thereon. A fluid passage extends between the inner chamber of the booster rocket and the shuttle glider support structure such that a portion of gas within the inner chamber is redirected from the inner chamber through the fluid passage and out of the shuttle glider support structure. The shuttle launch system further includes a shuttle glider having a fuselage and at least one wing, the shuttle glider removably mounted on the shuttle glider support structure on the booster rocket. Finally, the shuttle glider is separated from the booster rocket upon ejection of matter from the ejection charge of a model rocket engine held within the engine receiving chamber, the ejected matter compressing gas within the inner chamber, a portion of that gas being redirected from the inner chamber through the fluid passage and outward from the shuttle glider support structure thereby propelling the shuttle glider outward from the booster rocket.

Ozone depletion in the plume of a solid-fuelled rocket

Abstract: . The local effects of the emission of a solid-fuelled rocket on the stratospheric ozone concentration have been investigated by photochemical model calculations. A one-dimensional horizontal model has been applied which calculates the trace gas composition at a single atmospheric altitude spatially resolved around the exhaust plume. Different cases were tested for the emissions of the Space Shuttle concerning the composition of the exhaust and the effects of heterogeneous reactions on atmospheric background aerosol. The strongest depletion of ozone is achieved when a high amount of the emitted chlorine is Cl 2 . If it is purely HCl, the effect is smallest, though in this case the heterogeneous reactions show their largest influence. From the results it may be estimated whether ozone depletion caused by rocket launches can be detected by satellite instruments. It appears that the chance of coincidental detection of such an event is rather small.

Hybrid helium heater pressurization system and electrical ignition system for pressure-fed hybrid rockets

Abstract: A system for pressurizing a liquid oxygen tank (20) in a pressure-fed hybrid rocket (100) provides pressure at a first end of the tank (20) so that liquid oxygen can exit the second end. A hybrid heater (50) creates the pressure by mixing gaseous oxygen with the solid fuel (51) in the hybrid heater (50) to create an extremely high temperature. Helium is then mixed with the combustion products via an inlet (60) and this hot, inert mixture enters the liquid oxygen tank (20) via a diffuser (70). A second hybrid heater (90) pressurizes the gaseous helium sphere (40) to reduce the pressure decay caused by the withdrawal of the helium. Two of the hybrid heaters (50 and 90) are fed gaseous oxygen from a sphere (30). A novel ignition system is used to ignite the hybrid heaters (50, 90, 250). One of the hybrid heaters (250) is used to ignite the hybrid motor (100) in the rocket (11).

Nuclear rocket feed system incorporating an auxiliary power cycle

Abstract: A nuclear rocket engine comprising a primary feed system (4) for pumping rocket propellant from a propellant source (10) to a nuclear reactor (6) and an auxiliary feed system (60) coupled to the primary feed system. The auxiliary feed system includes a space radiator (84) for discharging excess reactor heat into space and a motorgenerator (74) for creating electricity from the excess reactor heat. A recuperator (20) operates to heat liquid propellant before it enters the reactor and to retain heat within the system. The auxiliary power system can be configured into a high thrust mode (see FIG. 1) for withdrawing heat from the engine when the reactor is operating at full power, a low thrust mode (see FIG. 2) for throttling propellant flow and radiating heat from the engine during reactor shutdown and a zero thrust mode (see FIG. 3) for cooling the nuclear reactor and generating electricity for the rocket's auxiliary power requirements for the remainder of the mission.