Aerospace & Aviation

Effect of jet and free stream air flow on wing or fuselage of jet and ducted-fan vtol aircraft in transition

Flow visualization techniques in helium at mach 15 & 20 - surface streamlines, attached and separated afterbody flow & stagnation points

Although the meteoroid hazard has not been adequately evaluated for proper design of space- craft for missions through cislunar and inter- planetary space, the extent of meteoroid inter- actions on space exploration may be estimated. The available information about meteoroids and cosmic dust obtained from ground-based obser- vations and satellite measurements is reviewed. This data is presently limited to meteoroids and cosmic dust with heliocentric orbits intersecting the plane of the ecliptic at one astronomical unit and interacting with the earth-moon system. From such information the expectation of the number of impacts with interplanetary matter as a function of exposed spacecraft surface area, exposure time, and meteoroid mass is presented. An estimate of the hazard has been made using appropriate hypervelocity cratering criteria. Some results of other authors are compared with the expectation of meteoroid interaction effects herein presented. The probability of damage to spacecraft is apparently considerably different and a bit less than some earlier estimates.

In any mission description, the vehicles, the flight profiles, and the astrionics hardware to im- plement the mission are all tightly interwoven things. A final result evolves only after many iterations to the solution are made. This paper will describe one of these iterations in the Saturn C-5 Earth Orbit Rendezvous approach to the Manned Lunar Landing Program. Since the iter- ation to be described in an nthone, there exists some basis for the hope that the perturbation from the final solution is small. This paper is not concerned with the landing itself, but only with those operations leading to injection of the space craft into the lunar trans- fer trajectory. However, as is to be expected, it is the target conditions which set the pace for the overall operation. The entire operation must be sized to culminate at a time and place which places the lunar target in an attainable position. The procedure would call for a burst of activity lasting over a relatively short time as compared to the long and extensive preparations leading up to it. The activity must be aimed at the opening of the lunar "launch window". Figure 1 illustrates the variation in the velocity increment required to launch a vehicle into a lunar transfer trajectory from a 485 kilometer earth orbit. The minima are at irregularly spaced intervals and are a func- tion of the inclination of the lunar and the earth satellite planes and of the position of the moon in its orbit around the earth (i. e. , the day of the month). In an operations analysis these spacings will influence the number of vehicles on the launch pad (primary and back-up) , their state of readi- ness, the firing rate, and also the flight profile to be chosen. Whether it is decided to go by "con- necting" or by "tanking" mode, the objective must be to get the spacecraft in the launch ready state at the opening of one of these launch windows. It may be desired that the first vehicle be capable of remaining in a functionally capable state even after bridging one or more of the gaps between the windows. This consideration will influence the design of the vehicles as well as the operational modes to be designed into the flight control hard- ware. For example, a sleep switch may be de- sirable from the standpoint of savings in battery we ight.

At the beginning of the Mercury Program it was realized that the entire preparation of the man for piloting the spacecraft would have to be done by use of simulation techniques. It was also realized that no single all-inclusive flight simulator could provide the astronaut with practice in all phases of the mission in a sufficiently realistic environment. To determine the simulation needs, the Mercury-Atlas flight was broken down into areas of astronaut mission responsibility and into areas of physiological and psychological factors affecting the astronaut. A study was made of existing and proposed devices to determine requirements, suitability, feasibility, and availability of these devices. The simulators and devices finally selected are listed in Figure 1. Omitted from this list are several devices that are not actually simulators, such as mockups and visual devices} disorientation devices such as the one at Pensacola, and airplane 0 g familiarization flights. The listings in Figure 1 are not by order of importance or time availability but rather in groupings of fixed-based and dynamic or moving-based simulators.

Parabolic manned re-entry design comparisons are made between two classes of slender and blunt configurations capable of conventional horizontal and vertical earth landings. Aerodynamic modulation using flaps and control augmentation requirements are discussed in terms of their effect upon re-entry trajectories and landing footprints. Within the environment defined by the re-entry trajectories, the convective and radiative modes of heat transfer are compared, their differences discussed, and the problems associated with the configurations for planetary return missions indicated. Thermal protection systems are discussed and the factors influencing the selection of a given system for the re-entry environment are indicated. Comparison between reflective and absorptive systems are made and the merits of each for a given flight time indicated. A thermo-structural analysis is presented which shows the trade-oft between structural operating temperature and heat protection system thickness requirements. Total heat shield and structural weights are compared for the landing footprints presented and their variation with range shown. Finally, the areas of uncertainty in the aero-thermo structural design analysis are indicated for each class of configuration studied, and the weight differences between conventional horizontal landing and vertical descent configurations are given.

Atmosphere controls for advanced manned space vehicles will in some cases require techniques different from those selected as optimum for the Mercury environmental control system (ECS). Water conservation will be rigorous for vehicles which use a solar or nuclear auxiliary power unit (APU). For missions longer than a few hundred hours, CO2 will be removed in a regenerable process. Several promising regenerable CO2 removal systems are presented, including freeze-out. Regeneration of oxygen from C02 is discussed as the next step in atmosphere control advancement. Test work on CO2 hydrogenation by the Sabatier process is cited. The recommended makeup of ECS elements is presented as a function of mission duration.

Lunar and planetary soft landings by gas-filled balloons

The orgin and nature of the moon

Effects of variable-energy electrostatic rocket-engine ion exhaust beams that pertain to the design of high-vacuum condensers are discussed. Three ion-engine beam sources, using cesium and mercury as propellants, were operated at energy levels of 200 to 9000 electron volts and currents of 0.002 to 0.200 ampere. Five condenser geometries, having surface areas from 0.287 to 5.76 square meters, were tested. Values for a function of the accommodation and condensation coefficients, which appears in an existing theoretical analysis, were determined. Although exhibiting rather wide variations in some tests, the values were estimated at approximately 0.15 for cesium and 0.015 for mercury. A method of estimating condenser surface area requirements is given.

Self-evaluating state vector control with applications to flexible liquid-fueled aerial and space vehicles

Orbit and ephemeris of periodic comet oterma

Miniaturized telemeter transmitting set

Costs of the united states space program

Ball bearing performance in liquid hydrogen