Propellant Transfer Research Articles

Modeling of cryogenic heated-tube flow boiling experiments of hydrogen and helium with the Generalized Fluid System Simulation Program

Accurate modeling of cryogenic boiling heat transfer is vital for the development of extended-duration space missions. Such missions may require the transfer of cryogenic propellants from in-space storage depots or the cooling of nuclear reactors. Purdue University in collaboration with NASA has assembled a database of cryogenic flow boiling data points from steady-state heated-tube experiments dating back to 1959, which has been used to develop new flow boiling correlations specifically for cryogens. Computational models of several of these experiments have been constructed in the Generalized Fluid System Simulation Program (GFSSP), a network flow code developed at NASA’s Marshall Space Flight Center. The new Purdue-developed universal correlations cover the full boiling curve: onset of nucleate boiling, nucleate boiling, critical heat flux, and film boiling. These correlations have been coded into GFSSP user subroutines. The fluids modeled in this study are liquid hydrogen and liquid helium. Predictions of local wall temperature and pressure drop are presented and compared to the test data.

Demonstration of charge-hold-vent (CHV) and no-vent-fill (NVF) in a simulated propellent storage tank during tank-to-tank cryogen transfer in microgravity

The space exploration from a low earth orbit to a high earth orbit, then to Moon, Mars, and possibly asteroids and moons of other planets is one of the biggest challenges for scientists and engineers for the new millennium. The enabling of in-space cryogenic rocket engines and the Lower-Earth-Orbit (LEO) cryogenic fuel depots for these future manned and robotic space exploration missions begins with the technology development of advanced cryogenic thermal-fluid management systems for the propellant transfer line and storage tank system. One of the key thermal-fluid management operations is the chilldown and filling of the propellant storage tank in space. As a result, highly energy efficient, breakthrough concepts for quenching heat transfer to conserve and minimize the cryogen consumption during chilldown have become the focus of engineering research and development, especially for the deep-space mission to Mars. In this paper, we introduce such thermal transport concepts and demonstrate their feasibilities in space for cryogenic propellant storage tank chilldown and filling in a simulated space microgravity condition on board an aircraft flying a parabolic trajectory. In order to maximize the storage tank chilldown thermal efficiency for the least amount of required cryogen consumption, the breakthrough quenching heat transfer concepts developed include the combination of charge-hold-vent (CHV) and no-vent-hold (NVF). The completed flight experiments successfully demonstrated the feasibility of the concepts and discovered that spray cooling combined with hold and vent is more efficient than the pure spray cooling for storage tank chilldown in microgravity. In microgravity, the data shows that the CHV thermal efficiency can reach 39.5%. The CHV efficiency in microgravity is 6.9% lower than that in terrestrial gravity. We also found that pulsing the spray can increase CHV efficiency by 6.1% in microgravity.

Parametric analysis of the Charge-Hold-Vent method for cryogenic propellant tank chilldown

In the absence of external heat exchangers, the on-orbit transfer of cryogenic propellants requires the receiver tank to first be quenched to a sufficiently low energy state to allow for a continuous no-vent fill to avoid unnecessary venting of liquid. One proposed method for tank chilldown that minimizes the potential for venting liquid is the charge hold vent (CHV) method. CHV follows a cyclic process that gradually removes thermal energy from the receiver tank by injecting liquid with the vent valve closed and allowing the fluid and wall to reach near-thermal equilibrium before venting the superheated vapor. However, the CHV method must be optimized to minimize complexity, mass, and time. This paper presents a modular CHV analytical model used to quantify the number of cycles and propellant mass consumed based on first principles. The model is used to examine the effect of eight parameters: receiver tank material, volume, mass, maximum expected operating pressure, and initial pressure, liquid injection pressure and temperature, and the target temperature. The model is validated against the only two available CHV datasets. Based on results, the tank mass-to-volume ratio is the most important factor in determining the number of CHV cycles and thus degree of difficulty in tank chilldown. The model can easily be used for early-stage design, sizing, and analysis of cryogenic propellant transfer systems.

Modeling of cryogenic heated-tube flow boiling experiments of nitrogen and methane with Generalized Fluid System Simulation Program

Accurate modeling of cryogenic boiling heat transfer is vital for the development of extended-duration space missions. Such missions may require the transfer of cryogenic propellants from in-space storage depots or the cooling of nuclear reactors. Purdue University in collaboration with NASA has assembled a database of cryogenic flow boiling data points from heated-tube experiments dating back to 1959, which has been used to develop new flow boiling correlations specifically for cryogens. Computational models of several of these experiments have been constructed in the Generalized Fluid System Simulation Program (GFSSP), a network flow code developed at NASA’s Marshall Space Flight Center. The new Purdue-developed correlations cover the full boiling curve: onset of nucleate boiling, nucleate boiling, critical heat flux, and post-critical heat flux boiling. These correlations have been coded into GFSSP user subroutines. The fluids modeled are nitrogen and methane. Predictions of wall temperature are presented and compared to the test data.

Validation of Universal Cryogenic Flow Boiling Correlations in Thermal Desktop for Liquid Helium

Understanding two-phase cryogenic propellant behavior is key to enabling technologies for future spaceflight missions. Developing accurate models of two-phase flow phenomena, particularly in the current work, flow boiling in the heating configuration, is relevant to the propellant transfer process both in 1-g and microgravity. Currently there is a need for more accurate, direct cryogenic data anchored models for various boiling phenomena. Recently, universal correlations for cryogens flowing in heated tubes have been developed for a wide variety of fluids, thermodynamic conditions, and various regimes across the boiling curve, and have been patched to provide a smooth, continuous predictive curve. This paper describes implementation and validation of these correlations into Thermal Desktop to improve predictive performance, with a focus on liquid helium. Results from Thermal Desktop using both the built-in and new correlations are validated against a historical dataset of flow boiling experiments in the heating configuration using liquid helium. Based on results, the new correlations show a substantial improvement over the original built-in flow boiling correlations in Thermal Desktop in predicting the wall temperature as a function of preponderant parameters for this quantum fluid at temperatures greater than the lambda temperature, T λ = 2.17K.

A Continuous Flow Boiling Curve in the Heating Configuration Based on New Cryogenic Universal Correlations

To enable the design of future in-space cryogenic propellant vehicles such as Lunar and Martian ascent and descent stages, fuel depots, and nuclear thermal propulsion systems, high-accuracy models of various phases of the propellant transfer process are required. This paper focuses on modeling steady-state flow boiling through the transfer line that connects a propellant tank to an engine or customer receiver tank, which is required to set limits on the allowable heat flux into the line. Using the largest-ever collection of available cryogenic heated tube data, universal cryogenic flow boiling correlations were recently developed for various regimes of the boiling curve and their transition points. However, to model flow boiling in heated tubes, these individual correlations must be patched together to provide a continuous predictive curve of wall superheat as a function of preponderant parameters. This paper provides an overview of the individual flow boiling correlations along with the logic and rationale for patching the correlations together to produce a single continuous boiling curve. Resulting flow boiling curves are presented for different possible permutations of independent inlet and flow conditions for illustration.

Liquid mass gauging during propellant transfer in microgravity

Advancing the precision of low-gravity liquid propellant mass gauging is a roadmap challenge for NASA and a significant obstacle to expanding human presence in space. We report flight test data for an implementation of a low-gravity liquid propellant mass gauging technology that is both non-invasive and propellant agnostic. The TRIO payload experiment flew twice on the Blue Origin New Shepard suborbital rocket and validated computational models of liquid distribution in microgravity at three different fill levels. The Propellant Refueling and On-orbit Transfer Operations (PROTO) payload experiment also flew twice on New Shepard and acquired mass gauging data before, during, and after a tank drain and fill that simulated in-space propellant transfers. Propellant volume estimates were derived at 1 Hz using the Modal Propellant Gauging (MPG) system. Analytical, semi-empirical, and finite element (FE) model results for predicted mode frequencies are compared to flight data for mode frequencies across a range of fill levels during the simulated propellant transfer. Flight (0-g), ground (1-g) and FEM mode frequencies for both 0- and 1-g are well-predicted by the semi-empirical models for frequency dependence on tank fill level.

Liquid nitrogen wicking rate experiments for screen channel liquid acquisition devices

A screen channel liquid acquisition device (LAD) commonly employed in microgravity space applications uses a finely woven mesh of metal wires along with capillary forces to separate liquid and vapor phases in storage tanks used for propellant transfer. Successful implementation of LADs in flight propulsion systems, particularly for cryogenic systems, depends on analytical design models that are anchored to experimental data. Screens are characterized using parameters such as bubble point pressure, flow-through-screen pressure drop, and wicking rate. The last parameter is especially important to cryogenic LAD systems due to the high propensity for screens to experience evaporation dry out during a mission. This paper presents new experimental wicking rate data for 12 screens covering a wide range of weave types, mesh counts, and metal type with liquid nitrogen (LN2) as a working fluid using a continuous flow method. The horizontal wicking rate data is used to validate the room temperature fluid-based wicking rate model from Grebenyuk and Dreyer (2016) with the gravitational term neglected. Results are examined parametrically, with respect to the effect of wicking rate direction, dewar test pressure, fineness of the screen, shute to warp wire ratio, warp wire orientation, and using cryogenic vs. room temperature fluid. The main takeaway from the current work is that the effective wicking diameter from cryogenic testing lies within experimental uncertainty as wicking diameters from room temperature data for all screens, implying that the room temperature based-model from Grebenyuk and Dreyer (2016) can be used to compute wicking velocities for cryogenic nitrogen for most screen types.

Propulsion Alternatives for Mars Transportation Architectures

Recently, NASA has pushed for returning humans to the moon, with in-situ resource utilization being the key capability to provide sustainability. One of the potential future developments could be a propellant depot in lunar distant retrograde orbit. Using Aerojet Rocketdyne Mars mission architectures and University of Alabama in Huntsville Nuclear Thermal Propulsion (NTP) engine models, this research analyzed the impacts of using chemical and engines as well as the liquid-oxygen (LOX) Augmented Nuclear Thermal Rocket (LANTR) engines for these missions and compared their performances to the reference hydrogen-based NTP (H-NTP) engines all the while assuming a propellant depot at lunar distant retrograde orbit. For a human mission to Mars originating in the lunar distant retrograde parking orbit, the LANTR engines will offer better overall performance than H-NTP engines with a predicted 55.6% decrease in propellant volume, 39% decrease in vehicle dry mass, and 50% decrease in the number of aggregation launches. This is due to LANTR’s 22% higher specific impulse than conventional chemical propulsion systems, three times higher density than pure hydrogen, and 440% higher thrust than the baseline H-NTP engines. However, these benefits come at the cost of the propellant mass, which is 32.4% higher for the conjunction class mission and 106.7% higher for the opposition class mission than the baseline H-NTP system.

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.

Evaluation of different interface-capturing methods for cryogenic two-phase flows under microgravity

The distribution of the gas–liquid interface is crucial to the accurate calculation of the flow and heat transfer of in-orbit cryogenic propellants, for which the surface tension force overtakes the gravitational force. As an essential oxidant, liquid oxygen has a lower surface tension coefficient and viscosity than most room-temperature fluids, causing a greater possibility of interface instability and breakage. Conventional numerical methods have seldom been assessed in terms of cryogenic two-phase flows under microgravity, and commercial software cannot provide a consistent platform for the assessment. In this study, a unified code based on OpenFOAM has been developed for evaluating four interface-capturing methods for two-phase flows, namely, the algebraic volume of fluid (VoF), geometric VoF, coupled level set and VoF (CLSVoF), and density-scaled CLSVoF with a balanced force (CLSVoF-DSB) methods. The results indicate that the CLSVoF-DSB method is most accurate in predicting the interface motion, because it uses the level set function to represent the gas and liquid phases. The gas–liquid interface predicted by the CLSVoF-DSB method is the most stable because it adopts the scaling Heaviside function to weaken the effects of spurious currents and increases the stability. The numerical algorithm of the algebraic VoF method is the most simple, so it has the highest efficiency. The geometric VoF uses the isoface to locate the gas–liquid interface in a grid cell, so it can obtain the thinnest interface. In applications of liquid oxygen, the CLSVoF-DSB method should be used if the overall accuracy is required.

Effects of microgravity on cryogenic spray quenching of a simulated propellant storage tank with enhancement by surface coating and flow pulsing

In-space cryogenic propulsion will play a vital role in space exploration mission to the Moon and future mission to the Mars. The enabling of in-space cryogenic rocket engines and the Lower-Earth-Orbit (LEO) cryogenic fuel depots for these future manned and robotic space exploration missions begins with the technology development of advanced cryogenic thermal-fluid management systems for the propellant transfer line and storage tank system.One of the key thermal-fluid management operations is the chilldown of the propellant storage tank. As a result, highly energy efficient, breakthrough concepts for quenching heat transfer to conserve and minimize the cryogen consumption during chilldown have become the focus of engineering research and development, especially for the deep-space mission to Mars. In this paper, we introduce such thermal transport concepts and demonstrate their feasibility in space for cryogenic propellant storage tank chilldown in a simulated space microgravity condition on board an aircraft flying a parabolic trajectory. In order to maximize the storage tank chilldown thermal efficiency for the least amount of required cryogen consumption, the breakthrough quenching heat transfer concepts developed include the combination of cryogenic spray quenching cooling, Teflon thin-film coating on the simulated tank surface, and spray flow pulsing. The boiling heat transfer physics that supports the concepts is clearly explained. The completed flight experiments successfully demonstrated the feasibility of the concepts and discovered that spray cooling is the most efficient cooling method for the tank chilldown in microgravity. In microgravity, the data shows that the chilldown thermal efficiency can reach 30% with Teflon coating alone. However, Teflon coating together with flow pulsing was found to substantially enhance the chilldown efficiency further to reach the 50% goal.

Development of a surface cryogenic propellant transfer concept for Martian operations

NASA is currently evaluating architectures to support human missions to Mars. A multitude of concepts are being traded and associated sensitivities are being analyzed. Mars Ascent Vehicle (MAV) propellant supply is a key consideration. Mass constraints for the Mars descent system and in-space transportation present significant architectural challenges that may preclude landing a fully fueled MAV for crew use. In such a case, to ensure the ability of the crewed mission to meet its objectives, propellant should be supplied to the MAV prior to the arrival of the crewed mission from Earth. A concept to robotically transfer liquid oxygen from a separate storage tanker across the surface to the MAV is proposed. This concept makes use of an unpressurized rover and is optimized to maximize the amount of propellant conveyed per trip.

Cryogenic spray quenching of simulated propellant tank wall using coating and flow pulsing in microgravity

In-space cryogenic propulsion will play a vital role in NASA’s return to the Moon mission and future mission to Mars. The enabling of in-space cryogenic engines and cryogenic fuel depots for these future manned and robotic space exploration missions begins with the technology development of advanced cryogenic thermal-fluid management systems for the propellant transfer lines and storage system. Before single-phase liquid can flow to the engine or spacecraft receiver tank, the connecting transfer line and storage tank must first be chilled down to cryogenic temperatures. The most direct and simplest method to quench the line and the tank is to use the cold propellant itself that results in the requirement of minimizing propellant consumption during chilldown. In view of the needs stated above, a highly efficient thermal-fluid management technology must be developed to consume the minimum amount of cryogen during chilldown of a transfer line and a storage tank. In this paper, we suggest the use of the cryogenic spray for storage tank chilldown. We have successfully demonstrated its feasibility and high efficiency in a simulated space microgravity condition. In order to maximize the storage tank chilldown efficiency for the least amount of cryogen consumption, the technology adopted included cryogenic spray cooling, Teflon thin-film coating of the simulated tank surface, and spray flow pulsing. The completed flight experiments successfully demonstrated that spray cooling is the most efficient cooling method for the tank chilldown in microgravity. In microgravity, Teflon coating alone can improve the efficiency up to 72% and the efficiency can be improved up to 59% by flow pulsing alone. However, Teflon coating together with flow pulsing was found to substantially enhance the chilldown efficiency in microgravity for up to 113%.

An advance in transfer line chilldown heat transfer of cryogenic propellants in microgravity using microfilm coating for enabling deep space exploration

The extension of human space exploration from a low earth orbit to a high earth orbit, then to Moon, Mars, and possibly asteroids is NASA’s biggest challenge for the new millennium. Integral to this mission is the effective, sufficient, and reliable supply of cryogenic propellant fluids. Therefore, highly energy-efficient thermal-fluid management breakthrough concepts to conserve and minimize the cryogen consumption have become the focus of research and development, especially for the deep space mission to mars. Here we introduce such a concept and demonstrate its feasibility in parabolic flights under a simulated space microgravity condition. We show that by coating the inner surface of a cryogenic propellant transfer pipe with low-thermal conductivity microfilms, the quenching efficiency can be increased up to 176% over that of the traditional bare-surface pipe for the thermal management process of chilling down the transfer pipe. To put this into proper perspective, the much higher efficiency translates into a 65% savings in propellant consumption.

Multifidelity Space Mission Planning and Infrastructure Design Framework for Space Resource Logistics

To build a sustainable space transportation system for human space exploration, the design and deployment of space infrastructure, such as in situ resource utilization plants, in-orbit propellant depots, and on-orbit servicing platforms, are critical. The design analysis and trade studies for these space-infrastructure systems require the consideration of not only the design of the infrastructure elements themselves, but also their supporting systems (e.g., storage and power) and logistics transportation while exploring various architecture options (e.g., location and technology). This paper proposes a system-level space infrastructure and logistics design optimization framework to perform architecture trade studies. A new space-infrastructure logistics optimization problem formulation is proposed that considers the internal interactions of infrastructure subsystems and their external synergistic effects with space logistics simultaneously. Because the full-size version of this proposed problem formulation can be computationally prohibitive, a new multifidelity optimization formulation is developed by varying the granularity of the commodity-type definition over the space logistics network; this multifidelity formulation can find an approximate solution to the full-size problem computationally efficiently with little sacrifice in the solution quality. The proposed problem formulation and method are applied to the design of in situ resource utilization systems in a multimission lunar exploration campaign to demonstrate their values.

Liquid hydrogen line chilldown experiments at high Reynolds numbers. II. Analysis

This paper presents in-depth analysis of the recently conducted high Reynolds (Re) number liquid hydrogen (LH2) transfer line chilldown experiments at NASA Glenn Research Center. Vertically upward trickle and pulse flow chilldown tests were conducted over a wide range of LH2 liquid temperatures (20.3 K – 24.2 K) and mass flow rates (0.0012 – 0.036 kg/s) typical of a proposed cryogenic propellant depot. Two-phase flow mapping is used to correlate flow visualization with LH2 stream temperature profiles. Thermodynamic state diagrams, reverse boiling curves, and heat transfer coefficients are used to explain the evolution of the chill down process at various locations. Analysis reveals stark differences in behavior between skin and stream temperature measurements due to annular flow and bubbly flow dominated cooling. Analysis also shows that hydrogen chilldown at high Re numbers promotes better mixing and thus more liquid contact along the wall earlier on, in comparison to chilldown with other fluids like nitrogen where classical inverted annular flow boiling is incurred.

Multi-objective layout optimization for an orbital propellant depot

The overall layout optimization design of an orbital propellant depot involves the optimization of shape, size, and positions of propellant tanks in functional module and the optimization of positions of equipment in service module, with the aim of making the carrying capacity of propellant, dry/wet ratio, and mass properties meet the allowable values. To alleviate the difficulty in dealing with the overall optimization problems involving two modules of the orbital propellant depot, a step-by-step modeling and solving strategy is presented. Two multi-objective optimization mathematical models for the tanks in functional module (model I) and the equipment in service module (model II) are constructed separately, which are solved one after another. In the solution process of the two models, model I is solved firstly and the obtained optimization solution is transmitted to model II as a known condition. We mainly focus on the layout optimization of equipment in the service module and give a batch component assignment and layout integration optimization method. In the proposed method, all the components are grouped firstly according to the functional subsystem, and then the obtained component groups are sorted in descending order of their feature values. Finally, the sorted component groups are added into the service module one by one for both assignment optimization and layout optimization. The computational results of the case study show that the obtained Pareto solutions meet the given allowable values of carrying capacity of propellant, dry/wet ratio, and mass properties of the orbital propellant depot.

Liquid hydrogen sloshing in superheated vessels under microgravity

We consider the influence of superheated walls on the axial sloshing motion of liquid hydrogen in an upright circular cylinder. The sloshing motion is induced by the capillary driven reorientation of the free surface after a step reduction of gravity. The step reduction is achieved in a drop tower facility which provides 4.7 s of compensated gravity conditions (Drop Tower Bremen). We conducted 15 experiments with two vessel radii of 26.2 mm and 20.15 mm, and various wall temperature gradients from isothermal up to 101.4 K/m. The vessels were filled with hydrogen only (single species system), and operated at a pressure of approximately 100 kPa. The saturation temperature of hydrogen at this pressure is 20.7 K. Right after the step reduction of gravity, the perfectly wetting liquid rises at the superheated wall and overshoots the steady state position of the new equilibrium, thus causing an axial sloshing motion which persists throughout the whole experiments time. It is shown that the wall superheat reduces the initial rise velocity and decreases the period of oscillation. A single sided model is presented which computes the transient temperature distribution in the wall. This model allows us to compute the heat flux into the liquid film at the wall, and therefor to determine the evaporated mass. We present a relation between the evaporated mass and the sloshing half-periods. The results contribute to the understanding of the behavior of cryogenic liquids, such as hydrogen, oxygen, and methane, in launchers, spacecrafts and future orbital propellant depots.

Screen Compliance Experiments for Application of Liquid Acquisition Device in Space

The purpose of a liquid acquisition device (LAD) is to separate liquid and vapor phases inside a spacecraft propellant storage tank in the reduced gravity and microgravity conditions of space so that vapor-free liquid can be extracted to the transfer line. A popular type of LAD called a screen channel LAD or gallery arm, uses a fine porous screen and surface tension forces of the liquid to allow pure liquid to flow through the screen while blocking vapor penetration. To analyze, size, and optimize the design of LADs for future in-space propellant transfer systems, models and data are required for the four fundamental influential factors for LAD systems, including bubble point, flow-through-screen pressure drop, wicking rate, and screen compliance for a wide variety of screen meshes. While there is sporadic data available for three of these parameters, there is no published quantitative data for screen compliance. During the transient startup of propellant transfer, the liquid must be accelerated from rest to the steady flow demand velocity, which causes the screen to deform or comply, so compliance data is required for accurate transient LAD analyses; most design codes only consider steady state analysis. This paper presents screen compliance experiments on 14 different screens, examining the effects of fineness of mesh, open area, and screen metal type on compliance. A basic equation of state is also developed and validated against the data which can be easily integrated into any transient LAD flow code to model propellant transfer.