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Vacuum jacketed components on launch umbilical tower service arms at Launch Complex 39 - bayonet joints in cryogenic propellant and vent lines
Two cryogenic systems are described which will provide cooling for experiments to be flown on Spacelab 2 in the early 1980's. The first system cools a scanning infrared telescope by the transfer of cold helium gas from a separate superfluid helium storage dewar. The flexible design permits the helium storage dewar and transfer assembly to be designed independent of the infrared experiment. Where possible, modified commerical apparatus is used. The second cryogenic system utilizes a specially designed superfluid dewar in which a superfluid helium experiment chamber is immersed. Each dewar system employs a porous plug as a phase separator to hold the liquid helium within the dewar and provide cold gas to a vent line. To maintain the low vapor pressure of the superfluid, each system requires nearly continuous prelaunch vacuum pump service, and each will vent to space during the Spacelab 2 flight.
Behavior of Ullage Bubbles during Blowdown in Low-g Experiment (BUBBLE) is a cryogenic tank depressurization drop tower experiment currently being designed at NASA Glenn Research Center. There is a need for liquid level rise data while venting a cryogenic propellant tank below the liquid saturation pressure in a reduced gravity environment. When a cryogenic tank is vented to a vapor pressure below the liquid saturation pressure, bubbles become entrained in the liquid causing the bulk liquid-vapor interface to rise. In reduced gravity, bubble rise velocity is diminished leading to a larger liquid level rise compared to a 1g environment. The purpose of the experiment is to gain further understanding of tank depressurization fluid physics and obtain high-fidelity data for model validation in 1g and reduced gravity environments. On-orbit cryogenic propellant tanks operating at high fill levels must efficiently manage venting operations to avoid the risk of liquid entrainment in the vent line, which could lead to asymmetric control thruster loads, freezing and clogging, and loss of liquid propellant. Validated models could be used to design settling and venting profiles to reduce risk and increase efficiency for cryogenic storage and transfer operations.
When using superfluid helium in low gravity environments, porous plug phase separators are commonly used to vent boiloff gas while confining the bulk liquid to the tank. Invariably, there is a flow of superfluid film from the perimeter of the porous plug down the vent line. For the Soft X-ray Spectrometer onboard ASTRO-H (Hitomi), its approximately 30-liter helium supply has a lifetime requirement of more than 3 years. A nominal vent rate is estimated as ~ 30 ug/s, equivalent to ~ 0.7 mW heat load. It is therefore critical to suppress any film flow whose evaporation would not provide direct cooling of the remaining liquid helium. That is, the porous plug vent system must be designed to both minimize film flow and to ensure maximum extraction of latent heat from the film. The design goal for Hitomi is to reduce the film flow losses to <2 ug/s, corresponding to a loss of cooling capacity of <40 uW. The design adopts the same general design as implemented for Astro-E and E2, using a vent system composed of a porous plug, combined with an orifice, a heat exchanger, and knife-edge devices. In this paper, design, on-ground testing results and in-orbit performance are described.
Space Station and satellite reservicing will require the ability to vent gas on orbit from liquid supply or storage tanks and to gage liquid quantity under microgravity conditions. In zero gravity, (zero-g) the vortex vent is capable of venting gas from a tank of liquid containing gas randomly distributed as bubbles. The concept uses a spinning impeller to create centrifugal force inside a vortex tube within a tank. This creates a gas pocket and forces the liquid through a venturi and back into the tank. Gas is then vented from the gas pocket through a liquid detector and then out through an exhaust port. If the liquid detector senses liquid in the vent line, the fluid is directed to the low-pressure port on the venturi and is returned to the tank. The advantages of this system is that it has no rotating seals and is compatible with most corrosive and cryogenic fluids. A prototype was designed and built at the NASA Johnson Space Center and flown on the KC-135 zero-g aircraft. During these test flights, where microgravity conditions are obtained for up to 30 sec, the prototype demonstrated that less than 0.10 percent of the volume of fluid vented was liquid when the tank was half full of liquid. The pressure volume temperature (PVT) gaging system is used in conjunction with the vortex vent to calculate the amount of liquid remaining in a tank under microgravity conditions. The PVT gaging system is used in conjunction with the vortex vent to gage liquid quantity in zero or low gravity. The system consists of a gas compressor, accumulator, and temperature and pressure instrumentation. To measure the liquid in a tank a small amount of gas is vented from the tank to the compressor and compressed into the accumulator. Pressure and temperature in the tank and accumulator are measured before and after the gas transfer occurs. Knowing the total volume of the tank, the volume of the accumulator, the volume of the intermediate lines, and initial and final pressures and temperatures, the mass of the gas leaving the tank is equated to the mass of the gas entering the accumulator. The volume of liquid remaining in the tank is calculated using the ideal gas law.
This paper presents a theoretical study of the gasdynamic behavior of the system used to vent methane gas from an open-cycle cryogenic cooler carried on the Nimbus F satellite. The flow of this gas from the cryogen tank to the vent-line exit takes place at very low Reynolds number, and is strongly affected by variations in the tube cross-sectional area, substantial heat addition, and important frictional effects. All three of these factors play a role in determining the choking mass flow. Finite-difference calculations are used to establish the mass flow rates for a range of cryogen temperatures. Finite-difference calculations are also used to determine the flowfield in the plume that results when the vented gas exits from the spacecraft. The momentum-flux distributions within the plume are then used to estimate the torques applied to the spacecraft when the plume impinges on several surfaces near the exit point. These torque estimates compare favorably with flight data.
Testing was conducted to find a suitable filter to help separate helium gas from the Spallation Neutron Source’s mercury process loop. This filter is proposed for installation on the gas-liquid separator (GLS) vent line to prevent a column of liquid mercury from flowing outside of the process loop shielding in the service bay. The filter’s desired properties are to prevent the passage of liquid mercury, provide minimal resistance to helium flow, resist radiation, operate for the lifetime of the target, be replaceable, and have a reliable construction. Testing concluded that the Mott Corporation POU-10-S series of filters provided the best performance available. The filters could stop most liquid mercury passage while allowing helium passage at nominal pressures above the GLS in standard conditions. This result has several implications for the design of the GLS vent system.
Over the past year the Nuclear Thermal Rocket Element Environmental Simulator (NTREES) has been undergoing a significant upgrade beyond its initial configuration. The NTREES facility is designed to perform realistic non-nuclear testing of nuclear thermal rocket (NTR) fuel elements and fuel materials. Although the NTREES facility cannot mimic the neutron and gamma environment of an operating NTR, it can simulate the thermal hydraulic environment within an NTR fuel element to provide critical information on material performance and compatibility. The first phase of the upgrade activities which was completed in 2012 in part consisted of an extensive modification to the hydrogen system to permit computer controlled operations outside the building through the use of pneumatically operated variable position valves. This setup also allows the hydrogen flow rate to be increased to over 200 g/sec and reduced the operation complexity of the system. The second stage of modifications to NTREES which has just been completed expands the capabilities of the facility significantly. In particular, the previous 50 kW induction power supply has been replaced with a 1.2 MW unit which should allow more prototypical fuel element temperatures to be reached. The water cooling system was also upgraded to so as to be capable of removing 100% of the heat generated during. This new setup required that the NTREES vessel be raised onto a platform along with most of its associated gas and vent lines. In this arrangement, the induction heater and water systems are now located underneath the platform. In this new configuration, the 1.2 MW NTREES induction heater will be capable of testing fuel elements and fuel materials in flowing hydrogen at pressures up to 1000 psi at temperatures up to and beyond 3000 K and at near-prototypic reactor channel power densities. NTREES is also capable of testing potential fuel elements with a variety of propellants, including hydrogen with additives to inhibit corrosion of certain potential NTR fuel forms. Additional diagnostic upgrades included in the present NTREES set up include the addition of a gamma ray spectrometer located near the vent filter to detect uranium fuel particles exiting the fuel element in the propellant exhaust stream to provide additional information any material loss occurring during testing. Other aspects of the upgrade included reworking NTREES to reduce the operational complexity of the system despite the increased complexity of the induction heating system. To this end, many of the controls were consolidated on fewer panels. As part of this upgrade activity, the Safety Assessment (SA) and the Standard Operating Procedures (SOPs) for NTREES were extensively rewritten. The new 1.2 MW induction heater consists of three physical units consisting of a transformer, rectifier, and inverter. This multiunit arrangement facilitated increasing the flexibility of the induction heater by more easily allowing variable frequency operation. Frequency ranges between 20 and 60 kHz can be accommodated in the new induction heater allowing more representative power distributions to be generated within the test elements.
Nucleonic cryogenic quality meter design and development and fluid monitoring through hydrogen vent line on Saturn S-IVB stage under orbital conditions
Transient thermodynamic analyses have been made for the Kennedy Space Center (KSC) LOX loading system proposed for the space shuttle. The primary objective of these analyses is to predict the sensitivity of system performance to operating procedures, flow rate, temperature, and quality at the orbiter (ORB)/ground service equipment (GSE) interface. A further objective is to establish required system chill-down times. These analyses also include the prediction of transient thermodynamic conditions for the facility vent line, shuttle main engine (SSME) feed line entrance, ORB/ET interface, and the ET entrance (tank bottom).
The development of a Dewar system for handling liquid helium under weightless conditions is described. Porous plug designs for the prevention of superfluid creep out of the dewar through the vent line were evaluated. For the purpose of designing a neck to provide a transition from the cold cavity to the outside, the loads carried by the neck and equipment supports were studied. Temperature, pressure, and mass flow instrumentation for monitoring Dewar performance were also evaluated. In addition, multilayer blankets consisting of aluminized Mylar separated by Dacron net sheets were designed to insulate the pressure vessel. The dewar system is suggested for use with the star tracking telescope aboard the relativity satellite.
Attention is given to the design parameters of the LH2/LO2 system in a description of the development of the external tank for the Space Shuttle Main Propulsion System (MPS). Design and functional changes for the MPS are reviewed, including: improvement of the vibroacoustic environment through the adaptation of a two-stage pilot to the existing basic valve design, a redesign of the GH2 vent line to accommodate increased deflections and loads, and the minimization of ice formation through slide joint insulation. The anti-geyser splash plate for the LO2 system is outlined in terms of fuel flow parameters under varying conditions of temperature and pressure, and test results for individual components are presented.
The results of tests and studies to investigate the utilization of solid hydrogen for cooling of spacecraft instruments and sensors are presented. The results are presented in two sections; the first describing the tests in which an existing single stage solid cooler was filled and tested with solid hydrogen and the second which describes the analysis and design of a catalytic converter which will be tested in the vent line of the cooler.
A variety of payload integrator-furnished, Mission Peculiar Equipment (MPE) has been incorporated by the Spacelab Mission 1 payload complement in order to satisfy each payload developer's interface and service requirements. Attention is given to the design features of secondary structures, the NASA Experiment Power Branching Distributor and ESA Junction Box, module cabling and vent lines, and passive thermal protection. A major task to be fulfilled during payload integration is the satisfaction of all instrument and facility servicing requirements without mutual interference.
Because the venting of free hydrogen gas to the atmosphere presents an extremely hazardous situation, it was necessary to devise a means for safe, controlled venting of the shuttle external tank gaseous hydrogen during and after liquid hydrogen tank loading. Several design concepts that were considered initially were discarded as unfeasible because of vehicle weight restrictions, high cost, and because the proposed structure was itself deemed a hazard due to the vehicle's nonvertical launch trajectory. These design concepts are discussed. A design employing a support structure/access arm attached to the fixed service structure was finally selected. The various design problems resolved included vent arm disconnect/drop interference, minimizing refurbishment due to launch damage, disconnect reliability, vehicle movement tracking, minimizing vent line pressure drop, and the presence of other vehicle services at the same centralized supply area. Six launches have proven the system to be reliable, efficient, and of nearly zero refurbishment cost.
The infrared detectors on the Infrared Astronomical Satellite (IRAS), which was placed into a polar orbit in January 1983, are cooled to a temperature of less than 3 K by thermal coupling to the main cryogenic tank (MCT) containing superfluid helium. A porous plug built into the vent line entrance acts as a superfluid helium liquid/vapor separator in zero gravity. A description of the IRAS MCT flight porous plug is presented, and tests of the plug in situ are discussed, taking into account submerged plug tests, a restart test, and a cold vapor flow test. Aspects of flow rate determination in the case of an unavailability of flight flow rate data are also considered.
Design features of a tested catalytic converter for altering vented cryogenic parahydrogen used as a coolant on spacecraft into a para-ortho equilibrium for channeling to other cooling functions are described. The hydrogen is expected to be stored in either liquid or solid form. A high surface area Ni-on-Si catalyst was selected for tests at an operating pressure of 2 torr at a ratio of 1000 gr catalyst for a gr/sec hydrogen flow. Cylindrical and radial flow geometries were tried and measurements centered on the converter efficiencies at different operating temperatures when the converter was placed in the vent line of the H2 cooler. Efficiencies ranging from 10-100 percent were obtained for varying flow rates. Further testing is necessary to characterize the converter performance under a wider range of operating temperatures and environments.
Due to uneven load sharing and lagging inflation rates, the design of the Large Main Parachute (LMP) cluster, used to recover the Space Shuttle steel case Solid Rocket Boosters, had to be modified. The cause of the problem was excessive variation in effective porosity in the crown area of the LMP during first stage inflation. The design modification consisted of adding horizontal ribbons above the existing vent band to reduce the vent porosity and better control the position and attitude of the vent lines. Performance of modified LMP's since introduction indicates that the load sharing between the clustered chutes has been significantly improved.