Liquid Hydrogen Propellant Research Articles

Combined Cycle Nuclear System Architecture for Crewed Mars Spacecraft Propulsion and Power

Nuclear thermal propulsion can potentially reduce the time of flight and spacecraft system mass needed for human spaceflight beyond cislunar space. This nuclear propulsion system has comparable thrust capability to chemically impulsive systems, which at about twice the specific impulse, can double the delta-velocity (ΔV) for the same propellant mass. However, the canonical problem for nuclear propulsion has always been that its benefits are shadowed by low technology readiness of a complex system. This paper describes a combined cycle nuclear thermal rocket (CCNTR) system architecture for propulsion and electrical power that comprises a 42-MWt-capable nuclear reactor core to provide 9.4 kN thrust on demand at a specific impulse of 940 s. The liquid hydrogen propellant flow through the rocket chamber cools the reactor during burns, thereby producing thrust while concurrently rejecting waste heat to space. The reactor also produces up to 100 kWe power for the spacecraft, eliminating the need for solar power generation and averting challenges associated with restarting a cold reactor for propulsive burns. Radiators reject the waste heat from electrical power production. Earth-to-Mars orbital transfers less than 100 days appear feasible assuming 680,000 kg of liquid hydrogen propellant and a vehicle dry mass of 83,000 kg that includes the 13,000 kg CCNTR system. Together, these results suggest that a CCNTR could be most promising to enable crewed missions to Mars.

Comparative molecular dynamics study of liquid hydrogen annular jets in subcritical and supercritical environments

In this paper, the injection characteristics of a liquid hydrogen propellant jet under subcritical and supercritical operating conditions are comparatively studied at the molecular scale using molecular dynamics methods in the context of spacecraft propulsion systems. To control the temperature and pressure within the environment and nozzle respectively, two sub-computational domains are set up for separate modeling. The sub-computational domain merging process eliminates the broken molecular bonds by using the computational domain reduction in the x-direction. The jet velocity, density and other physical property distributions at different temperatures and pressures were comparatively studied. It is found that the rate of change of the radial density of the jet under supercritical conditions is small. Based on this, the surface area and volume of the liquid phase region of the jet were extracted and calculated. Compared to result of subcritical case, the volume of the liquid hydrogen jet in the supercritical environment is reduced by 1.9%, while the surface area is increase by 6.7%. To explain the variation in the surface area due to molecular diffusion, the molecular stress results revealed that the stress in the hydrogen jet in supercritical environment is only 88.9% of that in subcritical environment.

Test data analysis of the liquid hydrogen chilldown and fill transfer experiments on a flightweight aluminum tank

Future long duration manned missions beyond Low Earth Orbit will likely require efficient methods with which to transfer cryogenic propellant from a depot storage tank to a customer spacecraft. An essential component of the refueling process is the chilldown and fill of the customer receiver tank, which must be performed by minimizing consumed propellant and/or transfer time. This paper presents test data analysis of cryogenic liquid hydrogen propellant transfer tests conducted on a thin-walled flightweight aluminum tank at the NASA Glenn Research Center. Four tests were conducted to assess the feasibility of a rapid chilldown and fill method used to minimize total transfer time. Pressure, temperature, fill level, and flow rate data are presented to examine the two-phase flow boiling and heat transfer as the test rapidly evolves. Test results show that three of the four tests were successful at accomplishing a final fill level of >90% in less than five minutes.

Key Technologies, Systems, and Infrastructure Enabling the Commercialization and Human Settlement of Low Earth Orbit, the Moon, and Cislunar Space

More than 50 years have passed since 2001: A Space Odyssey debuted in April 1968. In the film, Dr. Heywood Floyd flies to a large artificial gravity space station orbiting Earth aboard a commercial space plane. He then embarks on a commuter flight to the Moon arriving there 25 h later. Today, in this the 52nd anniversary year of the Apollo 11 lunar landing, the images portrayed in 2001 still remain well beyond our capabilities. This article examines key technologies and systems (e.g., in situ resource utilization, fission power, advanced chemical and nuclear propulsion), and supporting orbital infrastructure (providing a propellant and cargo transfer function), that could be developed by industry for both NASA and future commercial ventures over the next 30 years, allowing the operational capabilities presented in 2001 to be achieved, although on a more Spartan scale. Lunar-derived propellants (LDPs) will be essential to developing a reusable lunar transportation system that can allow initial outposts to evolve into settlements supporting a variety of commercial activities. Deposits of icy regolith discovered at the lunar poles can supply the feedstock material needed to produce liquid oxygen (LO2) and liquid hydrogen (LH2) propellants. On the lunar nearside, near the equator, iron oxide-rich volcanic glass beads from vast pyroclastic deposits, together with mare regolith, can provide the feedstock materials to produce lunar-derived LO2 plus other important solar wind implanted (SWI) volatiles, including H2 and helium-3. Megawatt-class fission power systems will be essential for providing continuous “24/7” power to processing plants, human settlements and commercial enterprises that develop on the Moon and in orbit. Reusable lunar landing vehicles will provide cargo and passenger “orbit-to-surface” access and will also transport LDP to Space Transportation Nodes (STNs) located in lunar polar (LPO) and lunar equatorial orbits (LLO). Reusable space-based, lunar transfer vehicles (LTVs), operating between STNs in low Earth orbit (LEO), LLO, and LPO, and able to refuel with LDPs, offer unique mission capabilities, including short transit time crewed cargo transports. Even commuter flights similar to that portrayed in 2001 appear possible, allowing 1-way trip times to and from the Moon as short as 24 h. The performance of LTVs using both RL10B-2 chemical rockets and a variant of the nuclear thermal rocket (NTR), the LO2-Augmented NTR (LANTR), are examined and compared. If only 1% of the LDP obtained from icy regolith, volcanic glass, and SWI volatile deposits were available for use in lunar orbit, such a supply could support routine commuter flights to the Moon for many thousands of years. This article provides a look ahead at what might be possible in the not too distant future, quantifies the operational characteristics of key in-space and surface technologies and systems, and provides conceptual designs for the various architectural elements discussed.

Modeling the effects of dissolved helium pressurant on a liquid hydrogen rocket propellant tank

A model was developed using NASA’s Generalized Fluid System Simulation Program (GFSSP) for the self-pressurization of a liquid hydrogen propellant tank due to boil-off to determine the significance of mixture non-idealities. The GFSSP model compared the tank performance for the traditional model that assumes no helium pressurant dissolves into the liquid hydrogen propellant to an updated model that accounts for dissolved helium pressurant. Traditional NASA models have been unable to account for this dissolved helium due to a lack of fundamental property information. Recent measurements of parahydrogen-helium mixtures enabled the development of the first multi-phase Equation Of State (EOS) for parahydrogen-helium mixtures. The self-pressurization GFSSP model was run assuming that the liquid propellant was pure liquid hydrogen and assuming helium dissolved into the liquid utilizing the new helium-hydrogen EOS. The analysis shows that having dissolved helium in the propellant does not have a significant effect on the tank pressurization rate for typical tank conditions (-423 °F and 30 psia).

Numerical modeling of self-pressurization and pressure control by a thermodynamic vent system in a cryogenic tank

This paper presents a numerical model of a system-level test bed—the multipurpose hydrogen test bed (MHTB) using the Generalized Fluid System Simulation Program (GFSSP). MHTB is representative in size and shape of a space transportation vehicle liquid hydrogen propellant tank, and ground-based testing was performed at NASA Marshall Space Flight Center (MSFC) to generate data for cryogenic storage. GFSSP is a finite volume-based network flow analysis software developed at MSFC and used for thermofluid analysis of propulsion systems. GFSSP has been used to model the self-pressurization and ullage pressure control by the Thermodynamic Vent System (TVS). A TVS typically includes a Joule–Thompson (J–T) expansion device, a two-phase heat exchanger (HEX), and a mixing pump and liquid injector to extract thermal energy from the tank without significant loss of liquid propellant. For the MHTB tank, the HEX and liquid injector are combined into a vertical spray bar assembly. Two GFSSP models (Self-Pressurization and TVS) were separately developed and tested and then integrated to simulate the entire system. The Self-Pressurization model consists of multiple ullage nodes, a propellant node, and solid nodes; it computes the heat transfer through multilayer insulation blankets and calculates heat and mass transfer between the ullage and liquid propellant and the ullage and tank wall. A TVS model calculates the flow through a J–T valve, HEX, and spray and vent systems. Two models are integrated by exchanging data through User Subroutines of both models. Results of the integrated models have been compared with MHTB test data at a 50% fill level. Satisfactory comparison was observed between tests and numerical predictions.

Thermonuclear Operation Space Lift

The “Project Orion” small fission bomb propulsion concept proposed the one-stage launching of large payloads into low Earth orbit, but it was abandoned because of the radioactive fallout into Earth’s atmosphere. The idea is here revived by the replacement of the small fission bombs with pure deuterium–tritium fusion bombs, and the pusher plate of Project Orion with a large magnetic mirror. The ignition of the thermonuclear fusion reaction is done by the transient formation of keV super-explosives under the high pressure of a convergent shock wave launched into liquid hydrogen propellant by a conventional high explosive.

Mathematical modelling of a cryogenic engine

Cryogenic propulsion systems using liquid hydrogen and liquid oxygen propellant combination are used in satellite launch vehicles in view of the higher specific impulse (ISP). Due to extreme low temperature of cryogenic propellant and low density and explosive nature of LH2, the development of cryogenic propulsion system is very complex and time consuming. In ISRO an indigenous cryogenic upper stage powered by an engine developing a nominal thrust of 73.5 kN in vacuum with a propellant loading of 12.8 tonnes is developed and successfully flight tested for the first time in GSLV D5 flight on 5th January 2014. A mathematical model is developed for finalizing the engine start and shut off sequence, ensuring smooth and safe ignition, predicting performance of subsystems under transient phase, nominal and off nominal conditions, control systems parameter settings and for the performance estimation of engine in sea level and flight. This paper highlights the configuration and working of a cryogenic engine, mathematical modelling of the engine, applications of the model in the cryogenic engine development and comparison of predicted values with the test results.

Cryogenic Boil-Off Reduction System

A computational model of the cryogenic boil-off reduction system being developed by NASA as part of the Cryogenic Propellant Storage and Transfer technology maturation project has been applied to a range of propellant storage tanks sizes for high-performing in-space cryogenic propulsion applications. This effort focuses on the scaling of multi-layer insulation (MLI), cryocoolers, broad area cooling shields, radiators, solar arrays, and tanks for liquid hydrogen propellant storage tanks ranging from 2 to 10m in diameter. Component scaling equations were incorporated into the Cryogenic Analysis Tool, a spreadsheet-based tool used to perform system-level parametric studies. The primary addition to the evolution of this updated tool is the integration of a scaling method for reverse turbo-Brayton cycle cryocoolers, as well as the development and inclusion of Self-Supporting Multi-Layer Insulation. Mass, power, and sizing relationships are traded parametrically to establish the appropriate loiter period beyond which this boil-off reduction system application reduces mass. The projected benefit compares passive thermal control to active thermal control, where active thermal control is evaluated for reduced boil-off with a 90K shield, zero boil-off with a single heat interception stage at the tank wall, and zero boil-off with a second interception stage at a 90K shield. Parametric studies show a benefit over passive storage at loiter durations under one month, in addition to showing a benefit for two-stage zero boil-off in terms of reducing power and mass as compared to single stage zero boil-off. Furthermore, active cooling reduces the effect of varied multi-layer insulation performance, which, historically, has been shown to be significant.

Screen channel LAD bubble point tests in liquid hydrogen

This paper presents experimental results for the liquid hydrogen bubble point tests for liquid acquisition devices (LADs) operating in low gravity cryogenic propulsion systems. The purpose of the test was to investigate parameters that affect screen channel LAD performance in a low pressure liquid hydrogen (LH2) propellant tank and to demonstrate several ways to increase the LH2 bubble point pressure. Three fine mesh screen channel LAD samples were tested in LH2 over the range of 16.7 K < T < 21.1 K and 31.5 kPa < P < 155 kPa using gaseous helium and hydrogen as pressurant gases. Results show that bubble point pressure is affected by screen mesh type, liquid temperature and pressure, and type of pressurization gas. Higher bubble points are achieved by using a finer mesh screen and pressurizing and subcooling the liquid with gaseous helium. In addition, there is evidence that the screen pore is itself temperature dependent.

Acoustic characterisation of a rectangular rocket combustor with liquid oxygen and hydrogen propellants

High frequency combustion instabilities have plagued the development of liquid propellant rocket engines since their invention. Continuing research efforts aim to understand the mechanisms by which the oscillating combustion chamber pressure of self-sustaining combustion instabilities is driven. To this end, a rectangular combustor with acoustic forcing, designated ‘BKH’, was developed to study flame-acoustic interaction under conditions which are representative of real rocket engines. This article describes the acoustic characterisation of the BKH combustor using high frequency dynamic pressure measurements from the first hot-fire tests using liquid oxygen and ambient temperature or cryogenic hydrogen injectants. Analysis of the resulting high frequency pressure measurements shows excellent agreement with the predicted acoustic properties of the system which were calculated using finite element methods.

Evaluation of supercritical cryogen storage and transfer systems forfuture NASA missions

Conceptual designs of Space Transportation Vehicles (STV), and their orbital servicing facilities, that utilize supercritical, single phase, cryogenic propellant were established and compared with conventional subcritical, two phases, STV concepts. The analytical study was motivated by the desire to avoid fluid management problems associated with the storage, acquisition and transfer of subcritical liquid oxygen and hydrogen propellants in the low gravity environment of space. Although feasible, the supercritical concepts suffer from STV weight penalties and propellant resupply system power requirements which make the concepts impractical.