ALBERT

Description: On the basis of the radiator area required for rejecting cycle waste heat, Rankine vapor cycles are far superior to the basic Brayton gas cycle for space turbogenerating powerplants. The present analysis considers modifications of the basic Brayton cycle and compares the modified cycles to the basic cycle with radiator area as the criterion of merit. The results indicate that reductions in radiator area attainable by modifying the basic Brayton cycle are small, and thus the competitive position of gasturbine cycles relative to Rankine vapor cycles is unchanged.

Description: The performance of an area ratio 200 bell-shaped nozzle, an area ratio 25 bell-shaped nozzle, and an area ratio 8 conic nozzle on a JP-4 fuel and liquid-oxygen rocket engine has been determined. Tests were conducted using a nominal 4000-pound-thrust rocket in the Lewis 10- by 10-foot supersonic tunnel, which provided the altitude environment needed for fully expanded nozzle flow. The area ratio 200 nozzle had a vacuum thrust coefficient of 1.96, compared with 1.82 and 1.70 for the area ratio 25 and 8 nozzles, respectively. These values are approximately equal to those for theoretical frozen expansion. The measured value of vacuum specific impulse for the area ratio 200 nozzle was 317 seconds for a combustion-chamber characteristic velocity of 5200 feet per second. The vacuum-specific-impulse increase for the area-ratio increase from 8 to 200 was 46 seconds.

Description: Equations relating the critical temperature and ion current density for surface ionization of cesium on tungsten are derived for the cases of zero and finite electric fields at the ion-emitting surface. These equations are used to obtain a series of graphs that can be used to solve many problems relating to ion-rocket theoretical performance. The effect of operation at less than space-charge-limited current density and the effect of nonuniform propellant flux onto the ion-emitting surface are also treated.

Description: An investigation was made of the relative influence of turbine inlet temperature, radiator temperature, and turbine efficiency on radiator area for Rankine cycles with rubidium, potassium, and sodium as working fluids. It was determined that, whereas turbine inlet temperature and turbine efficiency have gross effects on radiator size for a given inlet temperature a considerable latitude in the selection of radiator temperature may be accepted with only minor effects on required radiator size. Also investigated was the influence on turbine efficiency and design of the factors that distinguish alkali-metal vapor turbines from conventional gas turbines. The turbine configuration was determined to be a function of the involved working fluids and rotor blade speed. For a given blade speed, the number of stages required for high turbine efficiency was found to vary directly with turbine specific work output, and therefore to vary in the ratio 5 to 2.5 to 1 for sodium, potassium, and rubidium, respectively. Lower blade speeds than employed in conventional gas turbines may be required to satisfy critical stress considerations resulting from the elevated temperatures involved and the criterion of long-duration reliability. This will increase the number of turbine stages necessary to obtain high turbine efficiency and consequently increase turbine weight. The question of moisture formation was discussed and a calculation was made to indicate the nature of the aerodynamic losses due to moisture content. Various means of reducing moisture content were considered, including mechanical removal, increased radiator temperature, inefficient expansion, superheat, and reheat. Sample calculations were made in most cases to indicate their comparative effectiveness and resultant penalty in required radiator area.

Description: An investigation was conducted to determine whether solid-propellant rocket motors could be ignited and destroyed by small-particle impacts at particle velocities up to a approximately 10,940 feet per second. Spheres ranging from 1/16 to 7/32 inch in diameter were fired into simulated rocket motors containing T-22 propellant over a range of ambient pressures from sea level to 0.12 inch of mercury absolute. Simulated cases of stainless steel, aluminum alloy, and laminated Fiberglas varied in thickness from 1/50 to 1/8 inch. Within the scope of this investigation, it was found that ignition and explosive destruction of simulated steel-case rocket motors could result from impacts by steel spheres at the lowest attainable pressure.

Description: A nuclear-rocket regenerative-cooling analysis was conducted over a range of reactor power of 46 to 1600 megawatts and is summarized herein. Although the propellant (hydrogen) is characterized by a large heat-sink capacity, an analysis of the local heat-flux capability of the coolant at the nozzle throat indicated that, for conventional values of system pressure drop, the cooling capability was inadequate to maintain a selected wall temperature of 1440 R. Several techniques for improving the cooling capability were discussed, for example, high pressure drop, high wall temperature, refractory wall coatings, thin highly conductive walls, and film cooling. In any specific design a combination of methods will probably be utilized to achieve successful cooling.

Description: The Research Summary is a bimonthly report of supporting research and development conducted at the Jet Propulsion Laboratory. This periodical is issued in three volumes. Volume I contains summaries of the work accomplished by the Space Sciences, Systems, Guidance and Control, and Telecommunications Divisions of the Laboratory. Volume II contains summaries of the work accomplished by the Physical Sciences, Engineering Mechanics, Engineering Facilities, and Propulsion Divisions. All work of a classified nature is contained in Volume Ill.

Description: The Research Summary is a bimonthly report of supporting research and development conducted at the Jet Propulsion Laboratory. This periodical is issued in three volumes. Volume I contains summaries of the work accomplished by the Space Sciences, Systems, Guidance and Control, and Telecommunications Divisions of the Laboratory. Volume II contains summaries of the work accomplished by the Physical Sciences, Engineering Mechanics, Engineering Facilities, and Propulsion Divisions. All work of a classified nature is contained in Volume Ill.

Description: The Research Summary is a bimonthly report of supporting research and development conducted at the Jet Propulsion Laboratory. This periodical is issued in three volumes. Volume I contains summaries of the work accomplished by the Space Sciences, Systems, Guidance and Control, and Telecommunications Divisions of the Laboratory. Volume II contains summaries of the work accomplished by the Physical Sciences, Engineering Mechanics, Engineering Facilities, and Propulsion Divisions. All work of a classified nature is contained in Volume Ill.

Description: An investigation was undertaken to determine the effect of chamber and propellant feed temperatures on the starting characteristics of hydrogen peroxide thrust chambers. Start delay times for two types of thrust chamber designs in the 1- to 24-pound-thrust range were obtained over a range of chamber and propellant feed temperatures from 30 to 100 F. Start delay times obtained during the first minute of catalyst bed life and again after 6 minutes of total accumulated running time are presented as a function of chamber and propellant feed temperatures. The initial cold-start delay time of the hydrogen peroxide thrust chambers investigated was approximately 0.150 second to attain 90 percent of steady-state chamber pressure at chamber and propellant feed temperatures of 70 F and above. Both thrust chamber designs could be started at chamber and propellant feed temperatures as low as 30 F; start delay times did, however, generally increase at low temperatures. When the chamber was at an elevated temperature from a preceding firing, the start delay time was reduced to approximately 0.050 second, indicating a marked effect of chamber temperature at constant propellant feed temperatures. Accumulated run time affected the starting characteristics only when both the chamber and propellant feed temperatures were at reduced levels.

Description: An investigation was conducted to obtain nozzle performance data with relatively large-scale models at pressure ratios as high as 120. Conical convergent-divergent nozzles with divergence angles alpha of 15, 25, and 29 deg. were each tested at area ratios of approximately 10, 25, and 40. Heated air (1200 F) was supplied at the nozzle inlet at pressures up to 145 pounds per square inch absolute and was exhausted into quiescent air at pressures as low as 1.2 pounds per square inch absolute. Thrust ratios for all nozzle configurations are presented over the range of pressure ratios attainable and were extrapolated when possible to design pressure ratio and beyond. Design thrust ratios decreased with increasing nozzle divergence angle according to the trend predicted by the (1 + cos alpha)/2 parameter. Decreasing the nozzle divergence angle resulted in sizable increases in thrust ratio for a given surface-area ratio (nozzle weight), particularly at low nozzle pressure ratios. Correlations of the nozzle static pressure at separation and of the average static pressure downstream of separation with various nozzle parameters permitted the calculation of thrust in the separated-flow region from unseparated static-pressure distributions. Thrust ratios calculated by this method agreed with measured values within about 1 percent.

Description: A study has been made of a process in which a solar heating cycle is combined with an electrostatic cycle for generating electrical power for space vehicle applications. The power unit, referred to as a thermoelectrostatic generator, is a thin film, solid dielectric capacitor alternately heated by solar radiation and cooled by radiant emission. The theory of operation to extract electrical power is presented. Results of an experiment to illustrate the principle are described. Estimates of the performance of this type of device in space in the vicinity of earth are included. Values of specific power of several kilowatts per kilogram of generator weight are calculated for such a device employing polyethylene terephthalate dielectric.

Description: A simulator study was made of presentations and control requirements for a manned astrovehicle employed in the interception of artificial satellites during the terminal phase of an orbital rendezvous of the satellite. The study was considered in terms of a manned interceptor having a home berth at a manned space station which is in circular orbit 500 miles above the earth. Interceptions were restricted to coplanar conditions and ranges of one-half mile. The results are believed to be generally applicable to conditions wherein the target is in the terminal area. Presentations which do not provide for attitude indications of the intercepting vehicle are not satisfactory. A direct-visual-observation presentation employing a screen, a projected star background, and a projected image of the target provided a feeling of "realism" in tracking and would be a satisfactory pilot-training aid. Use of longitudinal-translation and attitude controls alone is inadequate. Use of translation control alone, parallel and normal to the axis of the intercepting vehicle, is effective; the addition of attitude control enhanced the effectiveness, provided the control was used discriminately. Direct-visual-observation interceptions can be performed effectively without the aid of range and rate-of-closure-of-range meters at speeds up to approximately 50 feet per second; however, higher speeds of interception require the use of these instruments.

Description: Nozzle performance data were obtained with three "method-of-characteristics" nozzles and a 150 conical nozzle at pressure ratios up to 130. Each basic configuration was cut off and tested at expansion ratios of 25, 20, 15, and 10. Unheated dry air was used at nozzle inlet pressures up to 22,000 pounds per square foot absolute. Nozzle thrust data were extrapolated to infinite pressure ratio (zero discharge pressure). As much as 1-percent increase in thrust with no increase in nozzle surface area (weight), can be obtained by using a method-of-characteristics, nozzle instead of a 15 conical nozzle when operating with a nozzle expansion ratio of 25 and nozzle pressure ratios from 200 to infinity. Conversely, for the same thrust, reductions in nozzle divergent surface area in the order of 25 percent are possible. The thrust performance of the method-of-characteristics nozzle was not as good as that of the 150 conical nozzle when operating at pressure ratios considerably below design (below 100 for the expansion ratio 25 nozzles). Theoretical and measured nozzle momentum coefficients agreed within about 0.6 percent. This is the order of accuracy of both the measured and theoretical values.

Description: A successful flight test of a spin-stabilized 20-inch-diameter solid-propellant rocket motor having a propellant mass fraction of 0.92 has been made. The motor was fired at altitude after being boosted by a three-stage test vehicle. Analysis of the data indicates that a total impulse of 44,243 pound-second with a propellant specific impulse of approximately 185 was achieved over a total action time of about 12 seconds. These results are shown to be in excellent agreement with data from ground static firing tests of these motors. The spherical rocket motor with an 11-pound payload attained a velocity of 15,620 feet per second (m = 16.7) with an incremental velocity increase for the spherical motor stage of 12,120 feet per second.

Description: An experimental investigation of exhaust diffusers has been conducted to evaluate various methods of minimizing the overall pressure ratio (from chamber to ambient pressure) required to establish and maintain full expansion of the nozzle flow (altitude simulation). Exhaust-diffuser configurations investigated were (1) cylindrical diffusers, (2) diffusers with contraction, and (3) diffusers including a right-angle turn. Cylindrical diffusers were evaluated with primary nozzles of various area ratios and types, as well as two clustered configurations; the other diffusers were evaluated with individual nozzles of constant area ratio and varied type. Air was the working fluid, except for two check points obtained with JP-4 fuel and liquid-oxygen rocket engines and cylindrical diffusers. The minimum length-diameter ratio of cylindrical diffusers was about 6 for minimum pressure-ratio requirements. With cylindrical diffusers of adequate length, the pressure-ratio requirements were primarily a function of the ratio of diffuser to nozzle-throat areas and were essentially independent of primary-nozzle type (including two clustered configurations) or area ratio. The two check points obtained with rocket engines indicated the pressure-ratio requirements at given ratios of diffuser to nozzle-throat areas were lowered, as compared with the requirements with air, as a result of the reduced ratio of specific heats. The minimum length-diameter ratio of the contraction throat of convergent-divergent diffusers was also about 6 for minimum pressure-ratio requirements. With adequate contraction-throat length, the pressure-ratio requirements of such diffusers were appreciably below those of comparable cylindrical diffusers when used with conical and cutoff-isentropic nozzles, but not when used with a bell nozzle. Minimum pressure-ratio requirements of a diffuser including a simple long-radius right-angle turn at maximum diffuser area, obtained with the center of radius of the turn a minimum of 2 diffuser diameters downstream of the nozzle exit, were not appreciably above those of a comparable optimum cylindrical diffuser. A diffuser including a long-radius right-angle turn at a contraction minimum area had somewhat lower pressure-ratio requirements than the aforementioned simple turn.