Propellant Tank Research Articles

Numerical and Experimental Study on the Nonlinear Liquid Sloshing in Cassini Tank

Liquid sloshing within propellant tanks of space vehicles has been a major concern in aerospace engineering. The aim of the work in this paper is to develop a flexible computational framework with high precision to simulate three-dimensional large-amplitude liquid sloshing in Cassini tanks. The finite element method is adopted to solve the fluid equations of motion in an arbitrary Lagrangian–Eulerian (ALE) framework, where the characteristic-based split method is combined with a fractional step method for solving the control equations in which the ALE kinematic description is incorporated to track the free liquid surface flexibly and effectively for large-amplitude liquid sloshing in Cassini tanks. In addition, an experimental platform is set up to verify the reliability and effectiveness of the presented method. The numerical results obtained from computer programming by using the ALE finite element method proposed in this paper are compared with the experimental results, proving the model to be successful. The nonlinear phenomena of large-amplitude liquid sloshing, especially rotary sloshing (steady-state swirling and non-steady-state swirling), in Cassini tanks under horizontal harmonic excitation are investigated by both the ground physical experiments and numerical simulations.

On-orbit mission overview of the low power hall thruster propulsion system aboard venμs satellite

Venμs is a satellite launched in 2017, for super-spectral Earth imaging and Electric Propulsion System (EPS) demonstration. In this paper we overview EPS design and operation throughout all five mission phases, from open/closed-loop orbit control, through orbit descent (720 → 410 km), orbit maintenance under high drag environment, to orbit raising (410 → 560 km). The EPS consisted of two throttleable Hall thrusters, PPU, Propellant Management Assembly (PMA), and a 9 L propellant tank carrying 16 kg of Xenon. Both thrusters operated in the 300–550 W power range, generated a combined total impulse of 158.1 kN-sec and consumed all propellant. Two methods are described to compute the remaining propellant mass – “Bookkeeping” and “PTV” methods. The advantages and disadvantages of each method are discussed in light of the Venµs mission. Thruster performance was measured on-orbit using the “Orbit Determination” method and compared to laboratory experiments conducted on the ground with identical thrusters. The measured performance on-orbit was found to agree within the error bars with the performance measured on the ground. Lastly, we present several repeating events in which the propulsion system suffered from sudden beam-outs or ignition difficulties. We present the methods used to construct a solution and implement it.

Cryogenic Insulation-Towards Environmentally Friendly Polyurethane Foams.

Cryogenics is the science and technology of very low temperatures, typically below 120 K. The most common applications are liquified natural gas carriers, ground-based tanks, and propellant tanks for space launchers. A crucial aspect of cryogenic technology is effective insulation to minimise boil-off from storage tanks and prevent frost build-up. Rigid closed-cell foams are prominent in various applications, including cryogenic insulation, due to their balance between thermal and mechanical properties. Polyurethane (PU) foam is widely used for internal insulation in cryogenic tanks, providing durability under thermal shocks and operational loads. External insulation, used in liquified natural gas carriers and ground-based tanks, generally demands less compressive strength and can utilise lower-density foams. The evolution of cryogenic insulation materials has seen the incorporation of environmentally friendly blowing agents and bio-based polyols to enhance sustainability. Fourth-generation physical blowing agents, such as HFO-1233zd(E) and HFO-1336mzz(Z), offer low global warming potential and improved thermal conductivity. Additionally, bio-based polyols from renewable resources like different natural oils and recycled polyethylene terephthalate (PET) are being integrated into rigid PU foams, showing promising properties for cryogenic applications. Research continues to optimise these materials for better mechanical performance and environmental impact.

Simulation analysis of cryogenic mechanical properties of carbon fiber reinforced silicon-containing epoxy resin composite

The use of carbon fiber-reinforced resin matrix composites instead of metals to manufacture cryogenic propellant tanks for spacecrafts is a development trend in the world aerospace industry. Cryogenic mechanical properties of the composites should be investigated in detail due to that the ultra-low temperature environment may cause micro cracks in the composites, leading to propellant leakage. In the present study, cryogenic tensile properties of a carbon fiber reinforced silicon-containing epoxy resin aomposite are investigated using experimental and numerical simulation methods. A silicon-containing epoxy resin with excellent cryogenic mechanical properties is developed by introducing a synthesized organic silicon polymer epoxidized polysiloxane and Nano-SiO2 into bisphenol F type epoxy resin. The tensile strength and modulus of the silicon-containing epoxy resin at −180 °C are 202.63 MPa and 7.81 GPa, which are 16.71% and 3.03% higher than those of bisphenol F type epoxy resin. The tensile strength of the silicon-containing epoxy resin at −180 °C is increased by 107.7% compared to that at room temperature, and the modulus at −180 °C is nearly twice that at room temperature. The carbon fiber reinforced silicon-containing epoxy resin composite is prepared by vacuum injection molding. The finite element model for the representative volume element of the composite unidirectional plate is established. The random sequence expansion method is used to randomly distribute the carbon fibers and simulate the thermal residual stress, the elastic performance, and the damage of the composite at cryogenic environments. Through comparison, it is found that the simulation results are in good agreement with the experimental results. The simulation reliability for cryogenic mechanical properties of carbon fiber reinforced resin matrix composites is verified. It is expected to provide a reference for the analysis and evaluation of cryogenic mechanical properties of composite tanks.

Microscopic damage behavior in CFRP cross-ply laminates at cryogenic temperature

For the practical use of cryogenic propellant tanks made of CFRP laminates, experimental elucidation of the laminates’ microstructural damage propagation behavior at cryogenic temperatures is important. This paper presents a newly developed tensile test rig that enables in-situ observation of microscopic damage using an optical microscope at cryogenic temperatures using a Gifford–McMahon refrigerator. In-situ observations of microscopic damage under uniaxial tensile loading were done at room temperature (290 K, 17 °C) and at 30 K (−243 °C) on cross-ply thin-ply CFRP specimens of three kinds with different 90° layer thicknesses. The results demonstrated that the crack propagation behavior is independent of temperature, that the matrix crack density is higher, and that the onset of matrix crack initiation strain is lower at 30 K than at 290 K. Furthermore, the thermal strain within 90° layer at 30 K and at 290 K was estimated using finite element analyses (FEA). The FEA results suggest that the decrease in onset strain of matrix crack initiation at 30 K is mainly attributed to the increase in the thermal strains within the 90° layer.

Study of the thermal characteristics of propellant tanks in microgravity and its application to propellant gauging

The use of the thermal propellant gauging (TPG) method for propellant mass detection in aerospace tanks inevitably involves the continuous heating of the tank, and the understanding of the heat transfer mechanism of the tanks in a microgravity environment plays a guiding role in the implementation of the TPG. In this paper, the thermal characteristics of propellant tanks under microgravity are investigated by simulation and it is found that with the weakening of gravity, the heat transfer slows down and ‘heat concentration’ occurs in the vicinity of the heater. The effect of this property on the implementation of the TPG was then investigated by simulation, and it was found that in the microgravity environment, the accuracy of the TPG detection can be improved by adjusting the locations of the heaters and temperature sensors on the external side of the tank wall.

Coupled modeling and simulation of tank self-pressurization and thermal stratification

Numerical analysis of thermo-fluid dynamics for cryogenic propellant storage primarily consists of nodal and computational fluid dynamics (CFD) modeling approaches. While nodal approaches prioritize faster computation over fidelity, CFD approaches promise accuracy and fidelity at the expense of significant computational resources. This paper presents an efficient coupling methodology developed to facilitate the simulation of a long-term self-pressurization process of a cryogenic propellant tank as well as associated thermal stratification phenomenon under both normal and reduced gravity conditions. Key highlight of the methodology is the development of a coupling scheme based on a domain decomposition approach that separates the tank into two computational regions at the liquid-vapor interface. SINDA/FLUINT, the nodal code, is utilized to simulate the liquid region, while ANSYS Fluent, the CFD code, handles the vapor region. A data exchange algorithm was developed to compute required boundary conditions for each region using the local thermodynamic properties of assumed infinitely thin interface. The coupling between the nodal and CFD codes was demonstrated by simulating a self-pressurization process in a small size tank using hydrogen as the working fluid. A sensitivity study on the grid sizing and coupling time step was conducted to determine appropriate spatial and ensure a divergent-free explicit data exchange time step size. Using the coupling approach, two numerical case studies were performed to study tank self-pressurization by considering two initial hydrogen fill levels. The effectiveness and efficiency of the coupling methodology was assessed by comparing the temperature and pressure evolution results of the coupled simulation with those obtained solely from the CFD simulation. Results showed that the coupling scheme and data exchange algorithm were successfully implemented, and the coupled simulation agreed well with the CFD simulations. In addition, a stratification number was used to characterize the transient dynamic development of the thermal stratification. Lastly, the computational costs associated with both approaches were compared. Significant reductions in simulation time were achieved for both cases.

Composite structures with embedded fiber optic sensors: A smart propellant tank for future spacecraft applications

Modern spacecraft and launch vehicle design is more oriented towards reducing system-level design and assembly complexities. In order to maintain high overall system performance while reducing these complexities, the use of smart materials and smart structural components is a well-known practice and is currently of rising interest to space systems' designers. The paper discusses a concept of smart space structures, in particular, a carbon fiber composites structure embedded with Optical Fiber Sensors (OFS) for spacecraft and launch vehicle applications. This study highlights the operational requirements for such tank and the smart features enabled by the optical fiber sensors. For the latter aspect, a quantitative comparison between Fiber Bragg Grating sensors (FBGs) and Distributed Optical Fiber Sensors (DOFS) based on Optical Frequency Domain Reflectometry (OFDR) is presented to state their core performance parameters, such as the sensitivity, sensing range, dynamic measurement capability, and spatial resolution. The increased performance and reliability in harsh environments associated with fiber optic sensors come with a reduction in size, mass, and power consumption compared to the conventional electronic sensors. Optical fiber sensors embedded in carbon fiber structures have proven their capability in providing accurate real-time measurements of temperature and monitoring structural integrity while detecting precisely possible points of rupture and failure as discussed and demonstrated in the literature review. The applications of fiber optic sensing in smart propellant tanks may extend to detecting fluid leakage, also providing increased precision in propellant gauging through temperature mapping, and can be used in on-ground qualification, pre-flight testing, as well as in-orbit operation, condition, and structural health monitoring. The article presents a statement for an optimal FOS embedding approach in composite pressure vessels and discusses the related placement and orientation method for the fiber optic sensors, coupled with a one component simplified analytical stress-strain transfer model deriving the stress component along the maximum principal direction (i.e., σMaxPrincipal). The novel approach is believed to serve the optimal employment of embedded FOS in composite structures, e.g., pressure vessels and light-weight structures in spacecraft, among other applications. The simplified model is believed to pave the way for effective data interpretation and processing, utilizing the available limited computational resources on-board the spacecraft.

EVALUATION OF THE SCATTER OF LIQUID LAUNCH VEHICLE POGO OSCILLATION AMPLITUDES DUE TO THE INFLUENCE OF THE SCATTER OF INTERNAL FACTORS

Almost all liquid launch vehicle developers faced the problem of ensuring stability in relation to POGO oscillations. The level of POGO amplitudes oscillations of the launch vehicle can be significantly affected by the scatter of internal factors. The study aims to create a mathematical model that can determine the range of POGO amplitudes in liquid launch vehicles. This will be demonstrated through the example of the Dnipro launch vehicle, which is affected by a variety of internal factors that cause its POGO amplitudes to vary. We developed the non-linear non-stationary mathematical model of POGO oscillations of the prototype of the Dnipro space launch vehicle. The model is built by taking into account the two lower vibration modes of the LV structure, two lower oscillation modes of the oxidizer feedline, and the first oscillation mode of the fuel feedline of the propulsion system. Modeling of dynamic processes was conducted in a combination of four liquid rocket engines based on the schematic of the staged rocket engine. The simulation takes into account cavitation phenomena in the engine pumps and delay times in the gas generators’ chambers. We have developed a method for determining the scatter of the POGO oscillations caused by the action of internal factors, which is based on the use of the LP uniformly distributed sequences. As internal factors, the frequencies, decrements, and shapes of LV structural oscillation modes, the values of pressurization of the propellant tanks, and the engines’ specific thrust impulses were considered. Based on the results of the calculations, the dependence of the POGO amplitudes in two regions of LV instability was determined, and the lower and upper enveloping curves for the POGO amplitudes were constructed. It is shown that the maximum POGO amplitudes oscillations in the first region of instability lie in the range from 0.23 g to 0.72 g and in the second region of instability — from 0 g to 0.60 g. Variants of combinations of internal factors, which provided the largest and smallest values of POGO amplitudes, were analyzed. This made it possible to determine the internal factors, the scatter of which has the greatest effect on the POGO amplitudes scatter: frequency, decrement, shape coefficients of oscillations of the oxidizer feedlines and the LV 1st mode structural longitudinal oscillations in the payload cross-section.

Numerical simulation of gas-liquid interface in a blade type propellant tank

Abstract Currently, most of the satellite propulsion systems are mainly on chemical propulsion, and liquid propellant is the necessary power fuel for the propulsion system. The propellant tank of the satellite is a critical unit used to store and manage propellant. The hard core of the blade type tank is the Propellant Management Device (PMD), which serves to provide non entrained propellant for the propulser under specified flow and acceleration conditions during all phases of satellite operation. In this paper, the management performance of propellant under microgravity environment in the tank of a blade type is studied. Using the VOF model, the liquid management performance under microgravity environment in the tank is numerically simulated. The repositioning process is numerical simulated. The good management performance of the tank is verified, and the fluid distribution law is obtained, which can provide guidance for the design of the blade type tank.

A numerical study on thermal-structural characteristics of cryogenic common bulkhead propellant tanks for various tank lengths and core thicknesses

Common bulkhead (CBH) propellant tanks have the advantage that its structural and space efficiency are high. The tank is a form in which an oxidizer tank and a fuel tank are integrally combined, and the oxidizer and fuel are separated by the CBH. The CBH propellant tank in which two tanks are integrated can reduce unnecessary space between the two tanks. However, there are other difficult problems regarding effectively designing the CBH propellant tanks even though there are structural and space benefit. This paper is a study on the thermal-structural characteristics of CBH tanks containing LOX/LH2 propellant. In order to conduct the thermal-structural analysis, the thermal analysis model was constructed using 2-dimensional axisymmetric elements to reflect the shape features of the tank, and the structural analysis model was built using 3-dimensional elements. The thermal-structural analysis was performed on the length of the bottom tank and the foam core thickness of the CBH. In the structural analysis of the tank, when the thermal load was ignored and only the pressure load was applied, there was a difference in the buckling load factor depending on the thickness of the CBH foam core, but as the thermal load was taken into consideration, there was also a difference in the bottom tank length. The analysis results show that the thermal-structural characteristics of the CBH propellant tank were influenced by the length of the bottom tank as well as the thickness of the foam core.

Rankine Vortex Suppression Employing Shaped Drain Ports: Assessment Using a New Parameter

The phenomenon of vortexing is common in engineering applications but must be prevented due to its adverse effects on draining systems. Environmental disturbances often induce rotational motions in liquid propellants, leading to the formation of vortex air cores within the propellant tank. These air cores can impede the flow of propellants through the outlet, potentially compromising engine performance and jeopardizing mission success. The formation of air-core vortices can also induce cavitation problems in the feed pumps of propellant-draining systems. To address these challenges, researchers have designed innovative drain port geometries aimed at suppressing vortex formations. However, the lack of standardized metrics for evaluating vortexing hinders the comparison of findings across different studies. To overcome this issue, the present study formulates and proposes a novel parameter that integrates various parametric characterizations used in previous studies to assess vortexing. Upon analyzing the performance of different drain port shapes using this newly developed parameter, this study identifies cruciform drain ports with an aspect ratio of 3 as the most effective ones in suppressing vortexing. The current findings not only contribute to the data on vortex suppression strategies but also highlight the need for standardized metrics to evaluate vortexing.

Thermal Modeling of CryoFILL liquefaction tests

Proposed NASA missions to Moon and Mars involve producing cryogenic propellant in-situ to reduce launch mass and requirements. One technique for liquefaction of the gases produced through electrochemical processes is to circulate cold gaseous neon or helium through broad area cooling tubes attached to the outside of the propellant tanks. To determine the performance of this liquefaction process tests were conducted at NASA Glenn Research Center in a 2.1 cubic meter tank with a broad area cooling network (CryoFILL). A thermal/fluid model of the tank and its cooling loop is developed in Thermal Desktop. Details of the model and the model predictions and comparison to some of the experimental data from the CryoFILL liquefaction tests are presented here.

Development of test rig for optical diagnostics of cryogenic spray

Spray cooling is the primary method for conducting chilldown and fill of cryogenic propellant tanks. In a typical spray injection process, a liquid sheet/jet that exits the spray nozzle undergoes primary breakup by the development of surface instabilities. Droplets and ligaments generated after the primary breakup undergo secondary breakup to create a dispersion of droplets, which extract heat on impact with the tank walls. In the existing literature, there is limited data on the primary breakup of cryogenic sprays and their detailed visualization. Moreover, an insight into the spray characteristics such as primary breakup length and cone angle is vital to the development of computational models. In this investigation, optical diagnostics of cryogenic spray breakup and measurements of cryogenic spray characteristics have been conducted. Liquid nitrogen is the selected cryogen for the present analysis. To capture the transient nature of spray breakup after the injection, a novel shadowgraph technique is developed to photographically freeze the spray motion. Cryogenic spray characteristics such as cone angle and primary breakup length are obtained from the shadowgraph. In addition, droplet velocity measurement using Particle Image Velocimetry is explored. The comprehensive experimental dataset obtained herein not only provides insights into the mechanism of spray formation for cryogenic fluids but also helps in designing of spray cooling system for tank chilldown.

Liquid hydrogen tank chill and no-vent fill prediction using computational fluid dynamics

Cryogenic tank chill and fill is an important cryogenic fluid management technology that supports and enables many of NASA’s long-duration space missions. For no-vent fill, the receiver tank pressure remains below the supply tank pressure during the entire duration of the fill so that the tank does not require venting. This is especially advantageous for tank fill operations in low gravity where the position of the liquid is not always known and venting the tank may cause loss of propellant by venting liquid. In lieu of expensive tests conducted on-orbit, accurate computational models capable of predicting receiver tank pressure during cryogenic propellant tank fill may be used to reduce system and propellant mass as well as mission risk. However, these numerical models must be validated or anchored to test data. This study presents computational fluid dynamics (CFD) models with conjugate heat transfer that are used to predict a liquid hydrogen tank chill and no-vent fill ground test conducted at Lewis Research Center (now Glenn Research Center) in 1991. The specific test case chosen for model validation implemented an upward-facing jet near the bottom of a cylindrical 34 liter tank. CFD predictions are compared to experimental measurements of tank pressure, fill level, and wall temperature. The CFD results show reasonable agreement to the test data but overpredict the pressure collapse near the end of the fill despite the liquid jet penetrating the liquid-vapor interface.

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.

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.

Experimental investigation on the discharge and suction gas performance in the multilayer insulation microchannels

The use of cryogenic propellants is prioritized for powering Mars lander propulsion systems. To provide thermal protection for these cryogenic propellants, a combination of multilayer insulation (MLI) and foam is commonly employed. However, during the Mars lander ascent, space travel, and reentry, unique phenomena such as transient venting and suction of gas occur within the MLI, significantly compromising its insulation performance. In the current study, an experimental platform was developed to investigate the discharge and suction of rarefied gas in MLI microchannels. Based on the perforation MLI results, the residual gas pressure in the innermost layer reaches below 10 Pa after over 100 h of evacuation. Baked MLI exhibits favorable discharge gas performance, with an average pressure value (the sum of the interstitial pressure in MLI divided by the total number of layers) 30.85% lower than that of the original MLI. Evacuation rates are improved by 11.87% when the MLI is flushed with nitrogen (N2). The baking condition does not significantly affect suction gas properties. The thermal conductivity of residual helium (He) is much higher than that of air and N2. These findings offer valuable insights for the design of cryogenic propellant tanks, aiding in optimizing their insulation performance.

Numerical simulation and experiment of a blade management device in microgravity

Currently, most satellite propulsion systems are mainly chemical propulsion, and liquid propellant is the necessary power fuel for the propulsion system. The working environment of microgravity occurs when the satellite is in orbit. According to the working characteristics of the satellite, a propellant management device of vane type is designed. The PMD is composed of a vertical guide plate and a liquid accumulator, which is applied to the propellant tank to store and manage the propellant. In this paper, the microgravity gas-liquid management and control mechanism of a propellant management device of vane type is studied by numerical simulation and tower drop test. The results of numerical simulation and tower drop test are consistent, which verifies the good fluid management characteristics of this type of propellant management device, and provides references for its engineering application and the design of this type of structure.

Novel Hypergolic Green Fuels with Hydrogen Peroxide for Propulsion Systems

By combining the advantages of chemicals from two different classes, a series of catalytically promoted green hypergolic fuels named polyamine/alkanolamine-based hypergolics (PAHyp) with highly concentrated hydrogen peroxide (96%) as an oxidizer was developed. In this paper, a novel recipe based on N,N,N′N′-tetramethylethylenediamine (TMEDA) and N-methyldiethanolamine (MDEA), named PAHyp 1, is characterized. Samples with different volume proportions of TMEDA and MDEA catalyzed with 0.5–4 wt% of copper salts were prepared. It was demonstrated that, by adding low catalyst loadings of 1–2 wt%, ultrafast ignition (as low as 8 ms) can be measured. Fast ignition is important to avoid a hard start in the startup phase, and reduced catalyst loading is important to avoid loss of performance in terms of specific impulse. To evaluate the system’s performance, an orbital transfer maneuver of a geostationary satellite that burns monomethylhydrazine with nitrogen tetroxide was considered. It was demonstrated that when using green formulations based on TMEDA/MDEA, smaller propellant tanks are required (because of the higher density of the green propellants), though more propellant mass is required due to a slightly low specific impulse. Besides good performance and reliable ignition, good storability is crucial for long-term space applications. Remarkably, although TMEDA is sensitive to air, visual inspection, ignition tests, and Fourier-transformed infrared spectroscopy analysis revealed that formulations with volume proportions of at least 60% of MDEA stored in vials filled with air survived with no signs of degradation after 14 months. Finally, to meet the requirement of long-term missions of several years in space, a hypergolic tripropellant feed system was proposed.