Showing papers in "Journal of Spacecraft and Rockets in 1964"

Mixing in turbulent axially symmetric free jets

Abstract: An analysis to solve mixing problems involving momentum, energy, and mass transfer in turbulent axially symmetric compressible flow is presented. The method is based on the linearization of the conservation equations in the plane of the von Mises variables while retaining the essential nonlinear nature of the equations in the physical plane. The application of this method to a laminar flow problem has been shown to be in good agreement with a numerical solution of the boundary-layer equations obtained by Pai. The extension of the method to turbulent flow requires an expression for the eddy viscosity. A possible modification of Prandtl's formulation, suggested by Ferri, is investigated. When this expression for eddy viscosity is introduced, the application of the present method to compressible turbulent flow yields analytical results that are in good agreement with available experimental data.

Nutational stability of an axisymmetric body containing a rotor

Abstract: T7K)R many satellite and space-probe missions, full three-F axis attitude control is not required, and simple (passive) spin stabilization is sufficient to insure satisfactory system performance. In many applications, in fact, the need for stabilization is not dictated by primary mission requirements at all but arises indirectly, perhaps as a means of minimizing or removing thermal and antenna design problems. Spin control is ideal for this type of mission. In other caess, considerable improvement in system performance is afforded, e.g., in a high-altitude communication satellite, increased antenna gain can be obtained using an antenna with a toroidal radiation pattern. Fully passive techniques for three-axis attitude control have only limited application, and it is generally necessary to employ far more sophisticated, active stabilization methods such as mass expulsion or flywheel control systems. However, there is one active, but fairly simple, three-axis control technique that is a natural extension of spin stabilization. In this technique, control of the third axis (control about the spin axis) is achieved by "despinning" a portion of the spinning spacecraft. That is, instead of the vehicle being a single body, it is constructed of two bodies (I and II in Fig. 1) constrained to rotate about a common axis. Suppose that II is the despun portion. A motor (not shown) is included which can change the relative rate about OZ. If the motor is servo-controlled, say by a sensor mounted on one of the bodies, the rate of body II can be maintained at zero or can be varied to permit a given axis in II, normal to OZ, to track an external reference. Thus, in effect, the despun component is three-axis stabilized by means of a single-axis active controller. Control of spin-axis attitude, if required at all, would normally be by ground command to activate, for example, pulse jets, or, in the case of earth orbiters, a magnetic coil mounted on the spinning component. A wide range of configurations can be devised which use this basic two-body construction. One application is id the use of a small directional antenna as the despun portion of a spin-stabilized communications satellite. In this application, the spin axis would be maintained normal to the orbit plane, and the antenna controlled to point along local vertical, allowing maximum gain to be obtained. Another type of single-axis control can be provided if the major part of a spacecraft is despun, and a high-speed internal rotor is employed to provide the gyroscopic stiffness. In fact, this case can be considered a derivative of a three-flywheel system from which two wheels are removed and the third operated about a high bias rate instead of null. Because of the high speed of the internal rotor, spin-axis pointing can again be controlled by ground command. The two-body configuration does appear to offer definite advantages as a means of achieving simplified three-axis attitude control of spacecraft; for this reason, it has received, and is receiving, serious attention in the industry. Indeed, one spacecraft of this construction, OSO I (built by Ball Brothers for NASA), was launched over two years ago and has operated successfully in orbit. In this paper the nutational stability of the two-body system is considered, and the constraints that the requirement for stability places on the moments of inertia of the configuration are established. SYMBOLS: O X , , O Y , , O Z ARE A X E S FIXED IN BODY I

Structural optimization and design based on a reliability design criterion

Abstract: The classical approach to design using a safety factor to represent conservatism often results in unnecessary weight and cost. A rational basis for optimization studies can be obtained by relating safety factor to reliability. The required statistics of the safety factor can be generated by Monte-Carlo and linear perturbation methods. Application of the perturbation method is discussed herein for assignment of design allowables, selection of design materials, and sizing structures for minimum weight.

Investigation of magnetohydrodynamic flight control

Abstract: This report presents the results of an investigation into the use of a magnetic field in the nose region of a hypersonic vehicle as a means of flight control. The vehicle is assumed to he moving at speeds that cause shock-ionization of a significant portion of the shock layer. The altitude-speed regimes in which this ionized flow will strongly interact with a magnetic field are determined. The governing interaction parameter Q' is essentially based on the shock standoff distance. The magnetohydrodynamic control forces, which can be generated within these large interaction regimes (()' > 1), are computed. It is shown that the critical portion of a typical flight path (down to 30 km and 6 km/sec) for interplanetary reentry lies well within the MHD control regime.

Thermal radiation from the exhaust plume of an aluminized composite propellant rocket

Abstract: : A technique is developed for calculating rocket base heating and spacecraft heating environments due to particle radiation from a single nozzle rocket exhaust plume. The technique has proved successful when applied to a single nozzle exhausting into a rarefied atmosphere on the basis of comparison of predictions with experimental results. The analysis treats radiation from a cloud of particles as that from an equivalent radiating surface. Thus, the problem is reduced to the determination of the proper values of the apparent surface emissivity and the effective temperature. In defining the apparent emissivity of the particle plume, an analogy with neutron scattering for a cylindrical cloud is adopted which shows the apparent emissivity to be dependent on particle emissivity and cloud optical thickness. Since the plume is non-uniform in particle size, concentration, and temperature, certain averaging techniques are used to define values of optical thickness and temperature.

Influence of injectant properties for fluid injection thrust vector control

Abstract: A linearized model of fluid-injection thrust vector control is developed. The analysis provides a very simple expression for injectant effective specific impulse and clearly shows effects of injectant and propellant properties on performance. Results compare favorably with a body of gas and liquid injection data available in the open literature. Aerothermochemical aspects are examined by predicting performance of selected injectants in combination with a hypothetical rocket propellant and nozzle. These injectants fall into six classes: inert gases, inert liquids, reactive gases, dissociative liquids, reactive liquids, and liquid bipropellants. Results are discussed in detail.

Exploitation of inertial coupling in passive gravity-gradient-stabilized satellites

Abstract: Parameters selected by analysis of passive damping techniques to facilitate optimum performance for an inertially coupled gravity gradient stabilized satellite

Gemini rendezvous program

Abstract: One of the main purposes of the Gemini program is to establish techniques for the rendezvous and docking of space vehicles. Since it is expected that these techniques will be used in many other space programs, the Gemini rendezvous experiments have been planned to provide the basic information needed, and provision has been made for accomplishing rendezvous in several ways, each having special features to be explored in the program so that an optimized operational procedure can be established.

Electrical propulsion for control of stationary satellites

Abstract: The application of electric propulsion engines to attitude control and stationkeeping of 24-hour stationary satellites is analyzed and compared with the performance of contemporary cold gas, monopropellant, and bipropellant propulsion systems. Both a 500-pound spin-stabilized and a 1500-pound three-axis controlled satellite compatible with current boost vehicles are examined, and each type of propulsion system compared as a function of mission duration and maneuver requirements. Solar electric propulsion is shown to be superior to chemical propulsion for long term stationkeeping and three-axis attitude control of the larger satellite. Cold gas and chemical propulsion are superior for attitude control and provide strong competition for electric propulsion in the stationkeeping of the smaller spin-stabilized satellite. (auth)

Failure mechanics in dense tungsten alloy rocket nozzles

Abstract: In the study of dense tungsten alloys for uncooled nozzle service, two principal failure mechanisms were found: erosion by chemical reaction with the propellant combustion gases and cracking due to thermal stress. In laboratory tests of heated filaments, the major chemical attack was oxidation by CO2 and H^O; the presence of CO and H2 markedly reduced these oxidation reactions. These conclusions on the reaction mechanism were substantiated by the erosion observed in subscale motor firings. Thermal-stress cracking was also studied in motor firings; it occurred early in the test from tensile stresses at the outer surface. Thermal upsetting from compressive stresses at the inner surface was noted. The fracture resistance of cast tungsten was superior to powder-process tungsten, and "hot working" generally increased the tendency to crack. Large grain size was superior to small grain size.

Hall-current accelerator utilizing surface contact ionization

Abstract: the weight parameters, shown in Table 3, were used for the aeroprecursor system. Figure 9 suggests an answer to the question whether precursors are necessary if "adequate atmospheric information is available/ It is not difficult to postulate a future time period in which uncertainties in our knowledge of the Martian atmosphere are possibly reduced to the point where the atmosphere can be used for braking purposes without data augmentation. However, the desirability, as opposed to the necessity, of incorporating precursors should be considered. Uncertainties in the knowledge of the atmosphere will still remain, and the propulsive requirements to attain the prescribed orbit after the atmospheric flight are a strong function of these uncertainties. Significant increases in pay load can also result from reductions in these propulsive requirements, as may be inferred from Fig. 9.

1-kw arcjet-engine system-performance test

Abstract: A 1-kw-d.c. arcjet-engine system (with power conditioning equipment and a cryogenic hydrogen propellant storage unit) was tested at the NASA Lewis Research Center in 1962 at an environmental pressure of 8 X 10 ~ mm of mercury. This radiation-cooled arcjet engine was designed to operate at approximately 1000-sec specific impulse. Current, voltage, propellant flow rate, thrust, chamber pressure, and body temperatures were measured continuously. Engine efficiencies (10-30%) and specific impulses (600-1400 sec) were deduced from the data, but reliability of the thrust measurements was poor. The exhaust plume was photographed with high-speed infrared film and various filters to reveal the size, turning angle, and shape of various wavelength regions. Plume radiation was measured by a radiometer with identical filters to determine the spectral distribution.

Analysis of Movable Louvers for Temperature Control

Abstract: Movable shutters or louvers are being employed on several spacecraft for active thermal control. The present analysis of an infinite louver array led to seven equations; six of them form a simultaneous set (three linear integral equations and three linear algebraic equations) which describes the heat-transfer characteristics of the array, and the seventh gives its relative thermal performance in terms of an effective emissivity. The equations are solved numerically, and effective emissivity is plotted as a function of louver blade position for various values of the dimensionless parameters that appear in the governing heat-transfer equations. The results are compared with experimental results obtained for the Mariner II spacecraft louver system; agreement is good when the test data are corrected for power dissipation from auxiliary hardware and the end losses of the finite array. The remaining 4.5% discrepancy at the full-open position is attributed to specular (rather than the assumed diffuse) reflection from the blades.

Selecting Cooling Techniques for Liquid Rockets for Spacecraft

Abstract: A* = nozzle throat area, in. AC = combustion chamber cross-sectional area, in. Ae = nozzle exit plane area, in. F = thrust, Ib h = heat-transfer coefficient, Btu/hr-ft-°R /sp = specific impulse, Ibf/(Ibm/sec) // = total impulse, Ib-sec k = thermal conductivity, Btu/hr-ft-°F L* = characteristic length, in. Pc = chamber pressure, psia P i nj = propellant injection pressure, psia q/A = heat flux, Btu/hr-ft TI = inside radius, in., ft Tg = gas recovery temperature, °R Tw = wall temperature, °R T co = effective heat sink temperature, °R t — wall thickness, in. AT = temperature difference between gas and wall V = velocity, fps WfC — coolant flow WT = propellant plus coolant flow, Ib/sec 6 = time, sec e = exterior wall emissivity a = effective thermal diffusivity, in./sec 5 = char depth, in. a= Stefan-Boltzmann constant, Btu/ft-°R-hr

Ion propulsion system for stationary-satellite control

Abstract: The application of ion propulsion to the 3-axes attitude control and station keeping of stationary satellites is discussed. The mission constraints that affect the engine system design, such as velocity increments associated with vernier orbit corrections, magnitude of disturbance torques, and required thrusting directions are presented. An evaluation is made of the tradeoffs between such critical parameters as attitude and station-keeping accuracy, average power utilization, duty cycle, thrust level, and satellite mass and moments of inertia. A preliminary design is given for the ion engine attitude-control and station-keeping system. (A prototype system has been developed and laboratory tested.) Such system parameters as power level, thrust level, specific impulse, and weight are specified.

Factors influencing configuration and performance of multipurpose manned entry vehicles

Abstract: Multipurpose manned reentry vehicle capable of entry at circular to hyperbolic velocities - design study

Results of studies on a twin-gyro attitude control system for space vehicles

Abstract: Automatic closed loop and pilot operated twin-gyro systems are discussed as attitude stabilizers for large space vehicles

Studies of touchdown stability for lunar landing vehicles

Abstract: Touchdown stability tests for lunar landing vehicles - symmetrical & asymmetrical inelastic impacts

Thermal similitude studies

Abstract: Thermal scale modeling for space vehicles - thermally similar model derived from mathematical model of prototype

Lunar Landing and Long-Range Earth Re-Entry Guidance by Application of Perturbation Theory

Abstract: A guidance scheme based on linear perturbation theory has been investigated. An improved capability has been achieved by the proper choice of independent variable and by appropriate weighting of the guidance gains computed by linear theory. The capability of this scheme applied to the descent-to-hover phase of lunar landing is demonstrated for two different types of nominal trajectory: a constant-thrus t gravity turn maneuver, and a constant-thrust, constant-pitch-rate maneuver. To demonstrate the performance of this type of guidance scheme for atmosphere entry, it has been applied to the guidance of a vehicle entering the earth's atmosphere at parabolic velocity. Its capability is evaluated for entries from abort conditions, as well as for entries within the normal entry corridor, and effects of variations in life-drag ratio and atmospheric density are investigated. It is shown that for both lunar landing and atmosphere entry this guidance system, which uses a single nominal trajectory, and therefore requires minimum storage capacity, permits guidance to a selected landing site from a wide range of initial conditions.

Analysis of self-cooling with infiltrated porous tungsten composites

Abstract: With the advent of solid rocket engines utilizing high-energy aluminized propellants, the need for a noneroding nozzle-throat insert material has become acute. The wall temperatures attained at the throat are above even the maximum practical operating temperature of tungsten. However, by infiltrating a porous tungsten insert with a second material, which is subject to gasification by boiling, sublimation, or decomposition, it is possible to reduce the wall temperature to a tolerable value. This paper presents an analysis of the behavior of this so-called "self-cooling" process with infiltrated porous tungsten composites. A one-dimensional, finite thickness, flat plate model is specified. As gasification of the infiltrant proceeds, the gaseous/liquid or solid infiltrant interface recedes from the surface. The transient partial differential equations describing the model consist of the continuity, momentum, and energy relations. The effect of transpiration is included. The equations are solved by a numerical finite-difference solution that was programed for the IBM 7094. Typical results of the computer program are presented for a number of organic and inorganic infiltrants at typical high-pressure solid rocket nozzle-throat conditions to illustrate the effects of certain variables such as porosity and permeability of the porous tungsten.

Electric space cruiser for high-energy missions

Abstract: Systems description of a spacecraft to be powered by a 500 kw nuclear turboelectric power plant for high energy missions

Freezing of liquids on sudden exposure to vacuum

Abstract: Evaporative cooling and freezing of liquids in a vacuum is analyzed. Preliminary consideration of the mechanism of the process resulted in a complex mathematical model. Therefore, it was necessary to make a number of simplifying assumptions to establish material and energy balances which were used, along with the kinetic theory of gases, as a basis for calculations. Results given for several cryogens and water indicate the effects of initial temperature, drop diameter, and ambient pressure on the time required for cooling to the triple point and subsequent freezing.