ALBERT

Description: A restraint system's main function is to restrain its occupant when his vehicle is subjected to acceleration. If the restraint system is rigid and well-fitting (to eliminate slack) then it will transmit the vehicle acceleration to its occupant without modifying it in any way. Few present-day restraint systems are stiff enough to give this one-to-one transmission characteristic, and depending upon their dynamic characteristics and the nature of the vehicle's acceleration-time history, they will either magnify or attenuate the acceleration. Obviously an optimum restraint system will give maximum attenuation of an input acceleration. In the general case of an arbitrary acceleration input, a computer must be used to determine the optimum dynamic characteristics for the restraint system. Analytical solutions can be obtained for certain simple cases, however, and these cases are considered in this paper, after the concept of dynamic models of the human body is introduced. The paper concludes with a description of an analog computer specially developed for the Air Force to handle completely general mechanical restraint optimization programs of this type, where the acceleration input may be any arbitrary function of time.

Description: Environmental problems of space flight structures - part 2, meteoroid hazards

Description: The ultimate objective of this project is to land men on the surface of the moon and return the men safely to earth. The objective of this document is to define the design approaches and operational techniques for transporting a three-man crew to the moon and returning them to earth.

Description: An investigation was made to determine the landing-impact characteristics of a reentry vehicle having a multiple-air-bag load-alleviation system. A 1/16-scale dynamic model having four canted air bags was tested at flight-path angles of 90 degrees (vertical), 45 degrees, and 27 degrees for a parachute or paraglider vertical letdown velocity of 30 feet per second (full scale). Landings were made on concrete at attitudes ranging from -l5 degrees to 20 degrees. The friction coefficient between the model heat shield and the concrete was approximately 0.4. An aluminum diaphragm, designed to rupture at 10.8 pounds per square inch gage, was used to maintain initial pressure in the air bags for a short time period.

Description: The simulation demonstrated linear and gimbal motions of the capsule and a Gemini-Agena docking.

Description: The performance characteristics of a pre-formed elliptical parachute at altitudes between 200,000 and 100,000 feet were obtained by means of in-flight photography. The tests demonstrate that this type of parachute will open at altitudes of about 200,000 feet if conditions such as twisting of the suspension lines or draping of the suspension lines over the canopy do not occur. Drag-coefficient values between 0.6 and 0.8 were found to be reasonable for this type of parachute system in the altitude range between 200,000 and 100,000 feet.

Description: This paper briefly describes three modes for accomplishing the Mars landing mission and compares them on a gross basis to indicate their probable order of merit and to identify design requirements placed on the Mars-excursion module (MEM) by the choice of mode. The paper shows that a flyby-rendezvous mode requiring low weight in earth orbit requires the MEM to enter the Mars atmosphere at velocities ranging from 20,000 to 30,000 ft/sec. The MEM for the flyby-rendezvous mode is not covered in this paper but merits further study. The MEM for the other modes of mission accomplishment begins its active operational sequence in Mars orbit and need not be greatly influenced by the method of delivery to Mars orbit. Parametric studies of the entry problem for two vehicles typifying a ballistic-type and a lifting-body-type were conducted to identify the problems associated with design of a MEM to accommodate the extremes of Mars atmospheric density presently predicted. This brief study indicates that: (a) the presently predicted density extremes of the Mars atmosphere present no serious design problems for a MEM which can operate across the entire band of predicted densities; (b) details of operational requirements and mission objectives will control the choice of configuration rather than entry requirements; and (c) the ballistic-type MEM is lighter and simpler but has less operational flexibility than a high L/D MEM.

Description: The Apollo II Second Generation Lunar Exploration System includes the direct landing spacecraft which consists of cargo command module, service module, and landing module. The landing module is also capable of being used as the Lunar Landing Vehicle (LLV) for landing unmanned cargos consisting of shelter modules such as the Lunar Occupancy Payload and other cargo in support of lunar surface operations. High energy cryogenic propellants are utilized to permit direct landing, manned, or logistic missions with use of a single Saturn V class booster. In last year's studies, the LLV was configured for maximum payload and with consideration for the direct three-man landing and return mission. Light weight and low vehicle height above the lunar surface at touchdown were major objectives. Logistic cargos of more than 27,000 pounds landed on the Moon were achieved within the single Saturn V boost capability. For the manned mission, the lunar take-off weight was determined to be 28,000 pounds ready for the return-to-Earth portion of the mission. The command module utilized was an advanced light-weight design weighing 10,000 pounds including supporting subsystems. Cryogenic oxygen/hydrogen propulsion was again used for maximum propulsion efficiency. Study of the Lunar Occupancy Payload was also accomplished last year. This module was configured to serve as an early lunar shelter or outpost station or as a basic module of an integrated base module complex. Single and dual compartment versions, as well as special mission versions, were studied.

Description: An experimental investigation was made to determine the landing characteristics of a 1/8-scale dynamic model of a reentry vehicle using a passive landing system to alleviate the landing-impact loads. The passive landing system consisted of a flexible heat shield with a small section of aluminum honeycomb placed between the heat shield and the crew compartment at the point that would be the first to contact the landing surface. The model was landed on concrete and sand landing surfaces at parachute letdown velocities. The investigations simulated a vertical velocity of 30 ft/sec (full scale), horizontal velocities of 0, 15, 30, 40, and 50 ft/sec (full scale), and landing attitudes ranging from -30 degrees to 20 degrees. The model investigation indicated that stable landings could be made on a concrete surface at horizontal velocities up to about 30 ft/sec, but the stable landing-attitude range at these speeds was small. The aluminum honeycomb bottomed occasionally during landings on concrete. When bottoming did not occur, maximum normal and longitudinal accelerations at the center of gravity of the vehicle were approximately 50g and 30g, respectively.

Description: Buffet and flutter characteristics of Saturn Apollo mission were studied using a dynamically scaled model. The model was built around a central aluminum tube for scaled stiffness distribution and strength to resist loads imposed during testing. Styrofoam sections attached to the core provided the correct external contours. Lead weights were added for correct mass distribution. An electromagnetic shaker was used to excite the model in its flexible modes of vibration during portions of the test. The model was supported on a sting, mounted by leaf springs, cables and torsion bars. The support system provided for simulating the full scale rigid body pitch frequency with minimum restraint imposed on elastic deflections. Bending moments recorded by sensors on the aluminum tube. Several modified nose configurations were tested: The basic configuration was tested with and without a flow separator disk on the escape rocket motor, tests also were made with the escape tower and rocket motor removed completely. For the final test, the Apollo capsule was replaced with a Jupiter nose cone. The test program consisted of determining model response throughout the transonic speed range at angles of attack up to 6 degrees and measuring the aerodynamic damping over the same range for the basic model and the modified configurations. Signals from the model pickup were recorded on tape for later analysis. The data obtained were used to estimate bending moments that would be produced on the full-scale vehicle by aerodynamic forces due to buffeting.

Description: The investigation was conducted in the Langley spin tunnel. The tunnel is an atmospheric wind tunnel with a vertically rising airstream in the test section and a maximum airspeed of approximately 90 feet per second. For this investigation, the model was hand launched into the vertically rising airstream. At times the model, both with and without a drogue parachute, was launched gently with as little disturbance as possible to determine what motions of the spacecraft were self-excited. At other times, the spacecraft with pre-deployed drogue parachute was launched into various spinning motions to determine the effectiveness of the drogue parachute in terminating these spinning motions. During drogue-parachute deployment tests, the spacecraft was launched into various spinning and tumbling motions and the drogue parachute was deployed. The motions of the model were photographed with a motion-picture camera, and some of the film records were read to obtain typical time histories of the model motion. The angles of attack indicated in the time histories presented are believed to be accurate within +/-1 degree. The mass and dimensional characteristics of the dynamic model are believed to be measured to an accuracy of: +/-1 percent for the weight, +/-1 percent for z(sub cg)/d, +/-15 percent for x (sub cg), and +/-5 percent for the moments of inertia. The towline and bridle-line lengths were simulated to an accuracy of +/-1 foot full scale.

Description: An experimental investigation has been made of some lunar-landing characteristics of a 1/6-scale dynamic model of a landing module having multiple-leg landing-gear systems. Symmetric four-point and five-point systems and an asymmetric four-point system were investigated. The landing-gear legs were inverted tripod arrangements having a telescoping main strut which incorporated a yielding-metal strap for energy dissipation, hinged V-struts, and circular pads. The landing tests were made by launching a free model onto an impenetrable hard surface (concrete) and onto a powdered-pumice overlay of various depths. Landing motion and acceleration data were obtained for a range of touchdown speeds, touchdown speeds, touch attitudes, and landing-surface conditions. Symmetric four-point and five-point systems and an Maximum normal acceleration experienced at the module center of gravity during landings on hard surface or pumice was 2g (full-scale lunar value in terms of earth's gravity) over a wide range of touchdown conditions. Maximum angular acceleration experienced was 12-1/2 radians/sec(exp 2) and maximum longitudinal acceleration was 1-3/4 g. The module was very stable with all gear configurations during landings on hard surface (coefficient of friction, microns=0.4) at all conditions tested. Some overturn instability occurred during landings on powdered pumice (microns=0.7 to 1.0) depending upon flight path, pitch and yaw attitude, depth of pumice, surface topography, and landing-gear configuration. The effect of stability of roll attitude for the limited amount of roll-attitude landing data obtained was insignificant. Compared with the four-point system, the five-point system with equal maximum gear radius increased landing stability slightly and improved the static stability for subsequent lunar launch. A considerable increase in landing stability in the direction of motion was obtained with an asymmetric four-point gear having two pads offset to increase gear radius by 33 percent in the direction of horizontal flight.