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Explorer 46 was launched into Earth orbit in August 1972 to evaluate the effectiveness of using double-wall structures to protect against meteoroids. The data from the Meteoroid Bumper Experiment on Explorer 46 is reexamined and it is concluded that most of the impacts originated from an Earth orbiting population. The probable source of this orbiting population is solid rocket motors fired in Earth orbit.
What is currently known about the orbital debris flux is from a combination of ground based and in-space measurements. These measurements have revealed an increasing population with decreasing size. A summary of measurements is presented for the following sources: the North American Aerospace Defense Command Catalog, the Perimeter Acquisition and Attack Characterization System Radar, ground based optical telescopes, the Explorer 46 Meteoroid Bumper Experiment, spacecraft windows, and Solar Max surfaces.
The theoretical performance of a new space radiator concept, the gas particle radiator (GPR), is calculated. The GPR uses a gas containing emitting, submicron particles as the radiating media. A transparent window contains the gas particle mixture around the solid radiator emitting surface. A major advantage of the GPR is that large emissivity (e sub T is greater than or = 0.8) is achieved without the use of emissive coatings. A radiation heat transfer analysis shows that for a modest volume fraction (approx. 10(-4)) of submicron particles and gas thickness (approx. 1 cm) the emissivity, e sub T, is limited by the window transmittance. Besides determining the emissivity, the window is the critical element for making it possible for the GPR to have lower mass than a tube type radiator. The window acts as a bumper to provide meteoroid protection for the radiator wall. The GPR was compared to a proposed titanium wall, potassium heat pipe radiator. For both radiators operating at a power level of 1.01 MW at 775 K it was calculated that the GPR mass was 31 percent lower than the heat pipe radiator.
A numerical model is developed for the interaction of a cometary probe, such as Giotto, with its environment, i.e., dust particles and gas. The spacecraft was set on a course to pass the comet at a velocity of 69 km/sec, so a chance existed that a potential field would form around the spacecraft and block lower energy particles from reaching the spacecraft instruments. The motion of electrons and ions is traced as a function of time to examine the evolution of the electric field, electric potential and the total space charge distributions on the surface of the spacecraft and its environment. Account is taken of the density of the particles and gas molecules at various distances from the comet, the collision energies involved, and the Giotto geometry. A solution is defined for the Poisson equation to describe the evolution of the plasma around Giotto, including the effects of ion collisions with the Al bumper protecting the spacecraft. The simulation predicts formation of an ion wake behind Giotto and the evolution of a positive potential on the order of 10 V around the spacecraft, i.e., sufficient for a positive potential barrier near the surface of the spacecraft.
The generic design of a 20-in. diameter viewport for the space station modules is discussed. It should possess the capabilities of meteoroid/debris protection (with no metallic cover), redundancies in its meteoroid/debris protection, and pressure sealing systems. In addition, it should provide ease of change out for maintenance or repair. The design does not take into account the bumper-shield effect of the outermost panes in the meteoroid/debris analysis.
The ballistic limits of single sheet and double sheet structures made of 6061 T6 Aluminum of 1.8 mm and larger nominal thickness were investigated for projectiles of 1.5 mm diameter fired in the Vertical Gun Range Test Facility and NASA Ames Research Center. The hole diameters and sheet deformation behavior were studied for various ratios of sheet spacing to projectile diameter. The results indicate that for projectiles of less than 1.5 mm diameter the ballistic limit exceeds the nominal 10 km/sec orbital debris encounter velocity, if a single-sheet suit of 1.8 mm thickness is behind a single bumper sheet of 1 mm thickness spaced 12.5 mm apart.
Abrupt change in load/deflection response controllable and predictable. Structurally-tailorable, nonlinear, snap-through spring (STNSTS) exhibits controllable and predictable abrupt change in load/deflection response and based upon known phenomenon of snap-through structural response. Composed of pin-connected two-bar linkage which depicts combined tension/compression springs. As load applied to STNSTS, stiffness is function of internal spring and bending stiffness of pin-connected bars. As load increases, bars deform laterally until they collapse and snap through. Has application in passively-tailored rotor-blade flap, pitch, and lag response, to improve aerodynamic performance and stability characteristics of rotors; in aerodynamically- and aeroelastically-tailored wing spars and ribs, to produce tailored deformation state for improved effectiveness in maneuvering, aerodynamic performance, and stability characteristics; and in energy absorbers for automobile bumpers and aircraft land
NASA's Orbital Maneuvering Vehicle (OMV) is being designed to withstand a 10-year lifetime in polar and low earth orbits. A large percentage of OMV's lifetime will be spent operating in the vicinity of the Space Shuttle and Space Station or in storage at these manned locations. An extensive analysis has been performed to determine the effects of the meteoroid and orbital debris environments on OMV's external fuel tanks. A finite element model of OMV was constructed using NASTRAN and analyzed with the meteoroid and debris design analysis code BUMPER. The results show that the long design lifetime, and the ever increasing man-made orbital debris environment, will require the use of shielding over the external fuel tanks.
A system of heat pipe radiators has been designed to provide waste heat rejection for an inertial fusion powered spacecraft capable of manned missions to other planets. The radiators are arrays of unfinned, arterial heat pipes operating at 1500 and 900 K. Liquid metal coolant carries up to 8000 MW of waste heat through feed pipes from on-board components (laser drivers and coil shield). The radiators do not rely on armor for protection from micrometeoroid penetration. An armored radiator design for this application with a 99 percent survivability would have a specific mass of 0.06 to 0.11 kg/kW at 1500 K. Instead, a segmentation of heat pipes is used, and bumpers are utilized to protect the feed pipes. This design reduces the specific mass to 0.015 to 0.04 kg/kW for the coil shield radiator (1500 K) and 0.06 to 0.12 kg/kW for the laser driver radiator (900 K).
A great deal of experimentation and analysis was performed to quantify penetration thresholds of components which will experience orbital debris impacts. Penetration was found to depend upon mission specific parameters such as orbital altitude, inclination, and orientation of the component; and upon component specific parameters such as material, density and the geometry particular to its shielding. Experimental results are highly dependent upon shield configuration and cannot be extrapolated with confidence to alternate shield configurations. Also, current experimental capabilities are limited to velocities which only approach the lower limit of predicted orbital debris velocities. Therefore, prediction of the penetrating particle size for a particular component having a complex geometry remains highly uncertain. An approach is described which was developed to assess on-orbit survivability of the solar dynamic radiator due to micrometeoroid and space debris impacts. Preliminary analyses are presented to quantify the solar dynamic radiator survivability, and include the type of particle and particle population expected to defeat the radiator bumpering (i.e., penetrate a fluid flow tube). Results of preliminary hypervelocity impact testing performed on radiator panel samples (in the 6 to 7 km/sec velocity range) are also presented. Plans for further analyses and testing are discussed. These efforts are expected to lead to a radiator design which will perform to requirements over the expected lifetime.
The global optimization of protective structural designs for spacecraft subject to hypervelocity meteoroid and space debris impacts is presented. This nonlinear problem is first formulated for weight minimization of the space station core module configuration using the Nysmith impact predictor. Next, the equivalence and uniqueness of local and global optima is shown using properties of convexity. This analysis results in a new feasibility condition for this problem. The solution existence is then shown, followed by a comparison of optimization techniques. Finally, a sensitivity analysis is presented to determine the effects of variations in the systemic parameters on optimal design. The results show that global optimization of this problem is unique and may be achieved by a number of methods, provided the feasibility condition is satisfied. Furthermore, module structural design thicknesses and weight increase with increasing projectile velocity and diameter and decrease with increasing separation between bumper and wall for the Nysmith predictor.
Spacecraft designers have always been concerned about the effects of meteoroid impacts on mission safety. The engineering solution to this problem has generally been to erect a bumper or shield placed outboard from the spacecraft wall to disrupt/deflect the incoming projectiles. Spacecraft designers have a number of tools at their disposal to aid in the design process. These include hypervelocity impact testing, analytic impact predictors, and hydrodynamic codes. Analytic impact predictors generally provide the best quick-look estimate of design tradeoffs. The most complete way to determine the characteristics of an analytic impact predictor is through optimization of the protective structures design problem formulated with the predictor of interest. Space Station Freedom protective structures design insight is provided through the coupling of design/material requirements, hypervelocity impact phenomenology, meteoroid and space debris environment sensitivities, optimization techniques and operations research strategies, and mission scenarios. Major results are presented.
The paper outlines selected outputs from an ongoing Marshall Space Flight Center/US Army Corps of Engineers parametric study of the effects of meteoroid/space debris impacts on Space Station Freedom module wall protection structures. The advantages and limitations of HULL hydrocode computer simulation are discussed, and examples of meteoroid/ debris impacts on various wall structural designs are presented. Trends in the terminal effects of particle sizes, velocities, shapes, and impact angles upon structural design parameters (wall and bumper thickness and spacing) are summarized and depicted through selected simulation runs. Throughout the paper, proposed structural modifications to the baseline module protection structure will be examined which are capable of mitigating damage from particle impacts.
Results are presented from a series of 22 hypervelocity impact tests carried out on the thermal protection system (TPS) for the Shuttle Orbiter. Both coated and uncoated low-density (0.14 g/cu cm) LI-900 and high-density (0.35 g/cu cm) LI-2200 tiles were tested. The results are used to develop the penetration and damage correlations which can be used in meteoroid and debris hazard analyses for spacecraft with a ceramic tile TPS. It is shown that tile coatings act as a 'bumper' to fragment the impacting projectile, with thicker coating providing increased protection.
At about 14,800 km from the Comet Halley nucleus, on the inbound leg, at least six of the sensors onboard the Giotto spacecraft observed an unusual, brief (about 30 to 500 ms) event: the ion-mass spectrometer data show a brief flow of energetic (up to several hundred electron volts) plasma consisting of protons, water group, and heavier ions. The Johnstone plasma analyzer data show a short burst of plasma, while the dust impact detector system data show an impact event in four of its detectors. The magnetometer signature of the event shows two brief dips in the field. The sudden change in the spacecraft attitude and spin rate observed by the camera at that same time has been interpreted as the result of a large (5 mg or more) dust-particle impact on the front bumper shield of the spacecraft. In addition, at about the same time the spacecraft star-tracker suffered damage. The report combines direct measurements of the composition and dynamics of a dust-impact plasma cloud, the dust particle mass, and the location of the impact on the spacecraft. Analysis of the data indicate that the impacting particle was water or ice-bearing, possibly loosely compared, and was composed of one or more of: carbon, nitrogen, and silicon.
Device uses friction to dissipate kinetic energy. When moving mass pushes in one direction, it offers substantial friction. Pushed in opposite direction, it offers negligible friction. Built-in spring resets for another shock-absorption cycle. Used in industrial machinery, automobile bumpers and suspensions, and parachute lanyards.
Analysis of interaction occurring between space debris and orbiting structures is of great interest to the planning and survivability of space assets. Computer simulation of the impact events using hydrodynamic codes can provide some understanding of the processes but the problems involved with this fundamental approach are formidable. First, any realistic simulation is necessarily three-dimensional, e.g., the impact and breakup of a satellite. Second, the thickness of important components such as satellite skins or bumper shields are small with respect to the dimension of the structure as a whole, presenting severe zoning problems for codes. Thirdly, the debris cloud produced by the primary impact will yield many secondary impacts which will contribute to the damage and possible breakup of the structure. The problem was approached by choosing a relatively new computational technique that has virtues peculiar to space impacts. The method is called Smoothed Particle Hydrodynamics.
A new technique for the interpolation/extrapolation of engineering data is described. The technique easily allows for the incorporation of additional independent variables, and the most suitable data in the data base is automatically used for each prediction. The technique provides diagnostics for assessing the reliability of the prediction. Two sets of predictions made for known 5-degree-of-freedom, 15-parameter functions using the new technique produced an average coefficient of determination of 0.949. Here, the technique is applied to the prediction of damage to the Space Station from hypervelocity impact of space debris. A new set of impact data is presented for this purpose. Reasonable predictions for bumper damage were obtained, but predictions of pressure wall and multilayer insulation damage were poor.