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A model of the Consolidated Vultee Aircraft Corporation Skate 7 seaplane was tested in Langley tank no. 2. Presented without discussion in this paper are landing stability in smooth water, maximum normal accelerations occurring during rough-water landings, and take-off behavior in waves.
A model of a Consolidated Vultee Aircraft Corporation Skate 7 sea-plane:was tested in Langley tank no= 2. Resistance data, 'spray photographs, and underwater photographs,are given in this report without discussion.
The information contained in the presentation covers development work carried out under SDS projects. The intent of the effort is to find chrome(VI) free alternates to our current chromated processing solutions. The information presented reports progress in work aimed to replace our alkaline cleaner (Turco 4215), conversion coat (Iridite 14-2) and chrome (VI) compounds used in LOx tank hydrostatic proof test solution. To date we have found candidates for use in the proof test solution and alkaline cleaner. These candidates are in the final stages of testing. Lab data is reported in the presentation.
The X-33 was a suborbital reusable spaceplane demonstrator, in development from 1996 to early 2001. The intent of the demonstrator was to lower the risk of building and operating a full-scale reusable vehicle fleet. Reusable spaceplanes offered the potential to lower the cost of access to space by an order of magnitude, compared with conventional expendable launch vehicles. Although a cryogenic tank failure during testing ultimately led to the end of the effort, the X-33 team celebrated many successes during the development. This paper summarizes some of the accomplishments and milestones of this X-vehicle program, from the perspective of an engineer who was a member of the team throughout the development. X-33 Program accomplishments include rapid, flight hardware design, subsystem testing and fabrication, aerospike engine development and testing, Flight Operations Center and Operations Control Center ground systems design and construction, rapid Environmental Impact Statement NEPA process approval, Range development and flight plan approval for test flights, and full-scale system concept design and refinement. Lessons from the X-33 Program may have potential application to new RLV and other aerospace systems being developed a decade later.
The measured Total Hydrocarbon Content (THC) levels in liquid oxygen (LOX) systems at Stennis Space Center (SSC) have shown wide variations. Examples of these variations include the following: 1) differences between vendor-supplied THC values and those obtained using standard SSC analysis procedures; and 2) increasing THC values over time at an active SSC test stand in both storage and run vessels. A detailed analysis of LOX sampling techniques, analytical instrumentation, and sampling procedures will be presented. Additional data obtained on LOX system operations and LOX delivery trailer THC values during the past 12-24 months will also be discussed. Field test results showing THC levels and the distribution of the THC's in the test stand run tank, modified for THC analysis via dip tubes, will be presented.
NASA is embarked on a new era of space exploration that will lead to new capabilities, new destinations, and new discoveries by both human and robotic explorers. Today, the International Space Station (ISS) and robotic probes are yielding knowledge that will help make this exploration possible. NASA is developing both the Orion crew vehicle and the Space Launch System (SLS) (Figure 1), that will carry out a series of increasingly challenging missions leading to human exploration of Mars. This paper will discuss the development and progress on the SLS. The SLS architecture was designed to be safe, affordable, and sustainable. The current configuration is the result of literally thousands of trade studies involving cost, performance, mission requirements, and other metrics. The initial configuration of SLS, designated Block 1, will launch a minimum of 70 metric tons (mT) (154,324 pounds) into low Earth orbit - significantly greater capability than any current launch vehicle. It is designed to evolve to a capability of 130 mT (286,601 pounds) through the use of upgraded main engines, advanced boosters, and a new upper stage. With more payload mass and volume capability than any existing rocket, SLS offers mission planners larger payloads, faster trip times, simpler design, shorter design cycles, and greater opportunity for mission success. Since the program was officially created in fall 2011, it has made significant progress toward launch readiness in 2018. Every major element of SLS continued to make significant progress in 2015. Engineers fired Qualification Motor 1 (QM-1) in March 2015 to test the 5-segment motor, including new insulation, joint, and propellant grain designs. More than 70 major components of test article and flight hardware for the Core Stage have been manufactured. Seven test firings have been completed with an RS-25 engine under SLS operating conditions. The test article for the Interim Cryogenic Propulsion Stage (ICPS) has also been completed. Major work continues in 2016 as the program continues both flight and development RS-25 engine testing, begins welding test article and flight core stage tanks, completes stage adapter manufacturing, and test fires the second booster qualification motor. This paper will discuss the program's key accomplishments to date and the challenging work ahead for what will be the world's most capable launch vehicle.
As developments of local production of cryogenic fluids on the Lunar or Martian Surface progress, it is important to understand transient system responses to help with the balancing of process plant power and understanding system level operations. During the Cryogenic Fluid In-situ Liquefaction for Landers (CryoFILL) testing, a series of transient oxygen liquefaction tests were completed. These tests included varying liquefaction flow rate, environmental temperature, tank fill level, and effective cryocooler lift while allowing the tank pressure to respond to the input controls. An additional transient test was run at the 90% fill level to determine the impact of injecting the gaseous oxygen at the bottom of the tank, allowing the vapor to bubble up through the liquid. Tests were run in a cyclical nature varying one variable at a time. The control variable was set in a manner to increase tank pressure for a period of time and then subsequently changed in a manner to decrease the tank pressure back to its original value with multiple cycles run for all tests. Tank pressure and system temperature responses were tracked as a function of time with an emphasis on repeatability. Results indicate that of the four variables tested, the environmental temperature is the least important. As expected, the bubbling of the liquefaction gas significantly decreased the pressurization and depressurization rates in the tank at the 90% full level.
The Space Shuttle LWT is divided into zones and subzones. Zones are designated primarily to assist in determining the applicable specifications. A subzone (general Specification) is available for use when the location of the component is known but component design and weight are not well defined. When the location, weight, and mounting configuration of the component are known, specifications for appropriate subzone weight ranges are available. Along with the specifications are vibration, acoustic, shock, transportation, handling, and acceptance test requirements and procedures. A method of selecting applicable vibration, acoustic, and shock specifications is presented.
Experimental data from several test series are compared to an existing correlation that predicts the amount of pressurant gas mass required to expel liquid hydrogen from axisymmetric tanks. It was necessary to use an alternate definition of the tank equivalent diameter to accommodate thermal mass in the tank wall that is initially warm and to accommodate liquid residuals in the tank after expulsion is stopped. With this modification, the existing correlation predicted mass requirements to within 14 percent of experimental results. Revision of the correlation constants using a nonlinear least-squares fit of the current experimental data has a minor effect, thus supporting the validity of the original correlation's form, its fitted constants, and the alternate definition of the tank equivalent diameter.
Metallic pressure tanks used in space missions are inherently vulnerable to hypervelocity impacts from micrometeoroids and orbital debris; thereby knowledge of impact damage and its effect on the tank integrity is crucial to a spacecraft risk assessment. This paper describes tests that have been performed to assess the effects of hypervelocity impact (HVI) damage on Titanium alloy (Ti-6Al-4V) pressure vessels burst pressure and characteristics. The tests consisted of a pair of HVI impact tests on water-filled Ti-6Al-4V tanks (water being used as a surrogate to the actual propellant) and subsequent burst tests as well as a burst test on an undamaged control tank. The tanks were placed behind Aluminum (Al) shields and then each was impacted with a 7 km/s projectile. The resulting impact debris plumes partially penetrated the Ti-6Al-4V tank surfaces resulting in a distribution of craters. During the burst tests, the tank that failed at a lower burst pressure did appear to have the failure initiating at a crater site with observed spall cracks. A fracture mechanics analysis showed that the tanks failure at the impact location may have been due to a spall crack that formed upon impact of a fragmentation on the Titanium surface. This result was corroborated with a finite element analysis from calculated Von-Mises and hoop stresses.
The Fluid Acquisition and Resupply Experiment, (FARE) is a Shuttle middeck-mounted experiment to demonstrate techniques for handling liquids in zero gravity for operations such as refueling spacecraft in orbit. The first flight took place on STS 53 launched December 2, 1992. Eight tests were performed during the mission and the experiment achieved 100 percent mission success. The second flight will be on STS 57, scheduled for launch in June 1993. The objective of FARE I was to demonstrate techniques for controlling the position of the liquid and gas within a tank during refilling and to better understand the operation of screen-type surface tension devices used to drain tanks in zero gravity. Tests were performed to demonstrate tank refilling, low gravity propellant slosh, and expulsion efficiency of the screen device. Expulsion efficiencies of 97 percent - 98 percent were demonstrated under a variety of flowrates and accelerations. Final fill levels of 60 percent to 80 percent were achieved during the vented fill tests.
In order to support long duration cryogenic propellant storage, the Cryogenic Fluid Management (CFM) Project of the Exploration Technology Development Program (ETDP) is investigating the long duration storage propertie$ of liquid methane on the lunar surface. The Methane Lunar Surface Thermal Control (MLSTC) testing is using a tank of the approximate dimensions of the Altair ascent tanks inside of a vacuum chamber to simulate the environment in low earth orbit and on the lunar surface. The thermal performance testing of multilayer insulation (MLI) coupons that are fabricated identically to the tank applied insulation is necessary to understand the performance of the blankets and to be able to predict the performance of the insulation prior to testing. This coupon testing was completed in Cryostat-100 at the Cryogenics Test Laboratory. The results showed the properties of the insulation as a function of layer density, number of layers, and warm boundary temperature. These results aid in the understanding of the performance parameters o fMLI and help to complete the body of literature on the topic.
The results are described of an experimental program which studied the performance of various no-vent fill techniques for tank-to-tank liquid transfer. The tests were performed using a cryogen simulant (Freon-114) and a test bed consisting of a multiple tank/plumbing network that enabled studies of a variety of different inlet flow and active mixing regimes. Several results and conclusions were drawn from the 26 transfer experiments comprising the program. Most notable was the significant improvement in fill performance (i.e., minimized fill time and maximized fill fraction) with increased agitation of the liquid surface. Another was the close correlation between measured condensation rates and those predicted by recent theories which express condensation as a function of turbulent eddy effects on the liquid surface. In most cases, test data exhibited strong agreement with an analytical model which accounts for tank heat transfer and thermodynamics in a 1 g environment.
The results are described of an experimental program which studied the performance of various no-vent fill techniques for tank-to-tank liquid transfer. The tests were performed using a cryogen simulant (Freon-114) and a test bed consisting of a multiple tank/plumbing network that enabled studies of a variety of different inlet flow and active mixing regimes. Several results and conclusions were drawn from the 26 transfer experiments comprising the program. Most notable was the significant improvement in fill performance (i.e., minimized fill time and maximized fill fraction) with increased agitation of the liquid surface. Another was the close correlation between measured condensation rates and those predicted by recent theories which express condensation as a function of turbulent eddy effects on the liquid surface. In most cases, test data exhibited strong agreement with an analytical model which accounts for tank heat transfer and thermodynamics in a 1 g environment.
A cryogenic test article, the Generic Research Cryogenic Tank, was designed to qualitatively simulate the thermal response of transatmospheric vehicle fuel tanks exposed to the environment of hypersonic flight. One-dimensional and two-dimensional finite-difference thermal models were developed to simulate the thermal response and assist in the design of the Generic Research Cryogenic Tank. The one-dimensional thermal analysis determined the required insulation thickness to meet the thermal design criteria and located the purge jacket to eliminate the liquefaction of air. The two-dimensional thermal analysis predicted the temperature gradients developed within the pressure-vessel wall, estimated the cryogen boiloff, and showed the effects the ullage condition has on pressure-vessel temperatures. The degree of ullage mixing, location of the applied high-temperature profile, and the purge gas influence on insulation thermal conductivity had significant effects on the thermal behavior of the Generic Research Cryogenic Tank. In addition to analysis results, a description of the Generic Research Cryogenic Tank and the role it will play in future thermal structures and transatmospheric vehicle research at the NASA Dryden Flight Research Facility is presented.
Instrumented model drogue tests were conducted in a ship model towing tank. The purpose of the tests was to observe and measure deployment and drag characteristics of such shapes as parachutes, crossed vanes, and window shades which may be employed in conjunction with free drifting buoys. Both Froude and Reynolds scaling laws were applied while scaling to full scale relative velocities of from 0 to 0.2 knots. A weighted window shade drogue is recommended because of its performance, high drag coefficient, simplicity, and low cost. Detailed theoretical performance curves are presented for parachutes, crossed vanes, and window shade drogues. Theoretical estimates of depth locking accuracy and buoy-induced dynamic loads pertinent to window shade drogues are presented as a design aid. An example of a window shade drogue design is presented.
This Technical Memorandum (TM) describes the thermal modeling effort undertaken at Marshall Space Flight Center to support the Cryogenic Test Laboratory at Kennedy Space Center (KSC) for a study of insulation materials for cryogenic tanks in order to reduce propellant boiloff during long-term storage. The Generalized Fluid System Simulation program has been used to model boiloff in 1,000-L demonstration tanks built for testing the thermal performance of glass bubbles and perlite insulation. Numerical predictions of boiloff rate and ullage temperature have been compared with the measured data from the testing of demonstration tanks. A satisfactory comparison between measured and predicted data has been observed for both liquid nitrogen and hydrogen tests. Based on the experience gained with the modeling of the demonstration tanks, a numerical model of the liquid hydrogen storage tank at launch complex 39 at KSC was built. The predicted boiloff rate of hydrogen has been found to be in good agreement with observed field data. This TM describes three different models that have been developed during this period of study (March 2005 to June 2006), comparisons with test data, and results of parametric studies.
The destabilizing effect of the liquid on attitude nutation stability was determined from an extensive series of inorbit tests. The liquid slosh driving frequency ratio (rotor nutation frequency/rotor spin rate) was varied over the range of 0.58 to 0.70 for the tests by rotating the spacecraft antenna platform at different rates in inertial space. A rotor mounted accelerometer sensed the spacecraft nutation. The observed time constant for the nutation angle increase or decrease was corrected for the stabilizing contribution of the platform mounted pendulum dampers to yield the net destabilizing dedamping contribution from the liquid slosh. The in orbit tests show two unexpected maxima in the dedamping contribution at driving frequency ratios that vary with the propellant loading. The rotor nutation frequency at the maxima was about one-third of the lowest mode liquid slosh frequency given by ground test data for unspun tanks, and thus did not correspond to a simple resonance of the liquid. Ground tests with spinning systems produced the same maxima, but the phenomenon is not yet understood.