Space Engineering Activities at CRYOSPACE and AIR LIQUIDE
By Jerome Lacapere, Air Liquide, Sassenage, France, and Mathieu Gardette, Cryospace, Les Mureaux, France
Engineering activities relative to the cryogenic propellant tanks of the European space launch vehicles were first developed at Air Liquide DTA(Advanced Technology Division) for the Ariane 4 upper stage tank (also referred to as H10). Subsequent development occurred at CRYOSPACE, a joint venture between Air Liquide (55%) and EADS ST (45%), for the Ariane5 main stage (EPC) and upper stage (ESC) liquid hydrogen (LH2) tanks. Development of the Ariane 5 upper stage liquid oxygen (LOX) tank (ESCLOX) and the helium sub-system (SSHEL) was later performed at Air Liquide.
Engineers at Air Liquide developed in-house software in the 60s and 70s to compute and predict the thermo-hydraulic behavior of the propellants (liquid hydrogen as fuel and liquid oxygen as oxidizer) in their respective tanks. The software models have been periodically improved, and they remain dedicated engineering tools. During the past five years, CRYOSPACE and Air Liquide DTA have also been using and benchmarking FLUENT for their specific cryogenic applications.
 Diagram of the Ariane 5 rocket, showing the cryogenic storage tanks |
The main function of a stage tank is to thermally condition the propellants in order to feed them to the engine in well-controlled temperature and pressure ranges. Over the years, a specific expertise has been developed and maintained for modeling the thermodynamics inside these tanks. In particular, the simulations must:
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compute the heat fluxes entering the tank volume when it is subjected to a severe external environment (ground conditions on the launch pad or flight conditions during the ascent phase, for example)
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accurately predict the heating rate of the liquid propellant and especially the stratified temperature profile that develops at the top of the tank due to natural convection along the tank walls
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compute the gaseous mass flow rate needed to maintain theullage pressure (or that in the space above the liquid) in a specified range throughout the flight (before and while the liquidis draining)
All of these computations lead to a calculation of the thermal residuals, which correspond to the mass of cryogenic liquid that is not compliant with the engine specifications in terms of temperature and pressure. Other calculations are performed to compute the diphasic residuals. These are the small amount of cryogenic liquid remaining in the nearly empty tank at the end of the engine feeding process. When this point is reached, external perturbations and the very high flow rate of ingestion could cause bubble ingestion, which must be avoided. Both types of residuals can be predicted using complex 3D simulations.
Thermal Residuals
Thermal stratification
Before launch, but following the filling of the tanks, the cryogenic propellants are thermally stabilized into a saturated state. Since the tank pressure values are typically close to 1.1bar, the stabilized temperature values are 20K (-253°C) for LH2 and 90K (-183°C) for LOX. A few minutes before launch,the tanks are pressurized with helium and brought to the flight pressure values. The pressurization process is aimed at stabilizing the tank structures,which will have to withstand significant mechanical loads during the ascent phase, and at providing a sufficient net positive suction pressure (NPSP, or difference between the static pressure in the tank and the saturation pressure in the pump) to prevent any cavitation in the turbo pumps during the draining (engine boost) phase.
 LH2 saturation curve over a range of temperatures, and the pressure evolution prior to and during launch and ascent |
If the tank pressure is regulated at a constant value during the flight, the NPSP depends mainly on the pressure loss in the engine feed lines between the tank and the turbo-pump, and the temperature of the propellant that is fed to the engine. The minimum acceptable NPSP value corresponds to a maximum acceptable draining temperature. The mass of propellant with a temperature exceeding this upper limit is unburnable. Since a launch vehicle always needs to optimize the ratio of its used propellant mass to its loaded mass, the residual mass should be minimized. It is therefore important to limit the heating rate of the propellant by an adequate insulation design and,if thermal stratification occurs, to be able to predict the draining temperature and its evolution over the full propulsive phase.
 Thermal stratification predicted by a 2D axisymmetric model of the ESC LH2 tank during draining; warmed propellant starts to stratify at the top of the liquid volume, and a hot spot appears at the tank bottom in a zero velocity area |
One place where thermal stratification occurs is in the upper stage liquid hydrogen (ESC LH2) tank. FLUENT was chosen to simulate this problem because of its volume of fluid (VOF) model with heat transfer, and its unstructured mesh capability. The goals of the simulations included predictions of thermal stratification and feed temperatures at the tank outlet with an accuracy better than ±0.1K. When deemed acceptable, the computations were done with 2D axisymmetric geometries. However, the particular shape and location of the feed line (tank outlet) sometimes required the use of a180°3D problem domain, which made use of a symmetry plane. One of the main concerns was the optimization of the mesh. In addition to capturing the turbulent convective boundary layers and moving free surface, there were local areas where the fluid velocity was known to be relatively high (in the boundary layer and close to the tank outlet), and other areas that could be considered dead zones with near zero velocity. Additional complexity was due to the fact that almost all of the parameters were time-dependant and many, such as heat fluxes on the walls, longitudinal accelerations, and draining flow rates, were highly variable.
Liquid sub-cooling
Although the more frequent concern is the heating rate of propellant inside the tanks, there are also reverse situations in which problems result from excessive sub-cooling. For the EPC tanks in the main stage, the LOX and LH2 tanks share a common bulkhead, and a strong thermal coupling occurs at the wall interface. At the bottom of the LOX tank, heat is conducted toward theLH2 side, which serves as heat sink. Locally, the liquid oxygen is driven below its saturation temperature into a sub-cooled state. The main stage engine must operate within a given temperature range at the pump inlet, however, and this leads to a requirement for a minimum sub-cooled temperature. Part of the functional studies performed on the propellant tanks included modeling the thermodynamic evolution inside the tanks and verifying that the feed temperatures do not drop below the allowable range.
 2D axisymmetric model of the bottom of the EPC LOX tank, showing a sub-cooled (below 91°K) layer of LOX growing in the tank bottom |
A 2D axisymmetric simulation was performed for the bottom of the LOX tank in the region of the common bulkhead. In addition to the liquid volume, the surrounding metallic tank structure and tank insulation were meshed as well. This allowed the external thermal environment specifications to be imposed directly as boundary conditions.
Natural convection inside the liquid propellant was simulated along with conductive heat transfer in the structure walls and insulation during the transient simulation. The results were used to predict the temperatures in the sub-cooled region and to make sure that the temperatures at the pump inlet were within the necessary range.
Temperature gradients in the tank structure
The tank structure is made primarily of aluminum alloys, and another set of simulations was performed for analyzing temperature gradients. Using fine meshes for the solid wall material, one objective of the simulations was to accurately compute the conductive heat fluxes resulting from thermal gradients, which can represent a fair amount of the heat budget entering the tank. The simulations were also used to export temperature profiles for input to mechanical analysis models. For these calculations, the temperature boundary condition was typically imposed on the cryogenic side of the structure at the wall/propellant interface.
As an example, an axisymmetric simulation was performed for the insulated lower skirt and ring of the upper stage LH2 tank. The tank was completely full, and still on the ground. The steady-state simulation showed a 265K temperature gradient along the 700 mm long skirt that generates a steady heat flux of 9 kW toward the tank inside.
 2D axisymmetric model of the insulated lower skirt and ring of the ESC LH2 tank |
Liquid sloshing in the tanks
The behavior of the cryogenic fluids inside the tanks is complex for a number of reasons. First, external perturbations during the launch cause sloshing and rolling of the free surface of the liquid.As learned from many years of development and flight measurement processing, the sloshing amplitudes experienced by the propellant can become high enough to modify the heat and mass transfer at the liquid/gas interface and to perturb the thermo-dynamic equilibrium in both phases. Typically these conditions cause the tank pressure to drop and the liquid temperature to rise. Second, in the gaseous dome above the liquids in the tanks, there is some amount of non-condensable helium present (which is used as a pressurizer). While it is not enough to violate the assumption of pure vapor in the LH2 tanks, it is significant in the LOX tanks, so must be taken into account in numerical modeling. Third, in the near future, the upper stages of Ariane will be required to operate in ballistic mode, and to have engine reignition following this phase of operation. During ballistic flight,the vehicle is in a micro-gravity environment. The walls of the tanks are wetted because of the predominance of capillary forces(with Bond number about 1) and because of the specific wetting property of cryogenic fluids (with contact angles less than 5°).Because of the increased surface area, the cryogenic liquid under-goes increased heating by external heat fluxes and by heat and mass transfer with the gas phase.
To address these needs, the VOF calculations need to take into account heat and mass transfer at the liquid/gas interface. A special model has been developed at Air Liquide for this purpose,and is now undergoing a complete validation. As part of the validation, characteristic tests have been performed with cryogenic liquids, including sloshing tests in a small cryostat (diameter~ 20cm) with liquid nitrogen and liquid oxygen (performed in the framework of the French-German program COMPERE). These tests have shown the characteristic evolution of gaseous pressure and propellant temperature that are directly linked to the heat and mass transfer at the interface. In some cases, the pressure evolution can be large, and this was shown to result in significant condensation at the free surface. The numerical analysis of these experiments was carried out using FLUENT and the special heat and mass transfer model was incorporated through a user-defined function (UDF). The simulation results were found to be in good agreement with the experimental data.
 Temperature stratification in the cryostat during sloshing, with the range limited to 85K |
 Volume fraction of helium in the gaseous phase during sloshing, showing the condensation of nitrogen close to the free surface |
After a complete validation of the model, it will be used to simulate thermal stratification and pressure evolution in a cryogenic tank in a micro-gravity environment, so that temperature and pressure can be computed precisely for the ballistic phase.
Diphasic Residuals
As the tank draining nears the end, some free surface disturbances may occur, particularly for the LOX in the upper stage tank, leading to sudden gas inclusion in the outflowing liquid. Because feed pump operation is not recommended in such conditions, the engine must be stopped before any gas ingestion occurs, and a residual mass of liquid will remain in the tank. The quantity of diphasic residual mass depends on the dynamic conditions during the draining process.
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 Geometry of the LOX tank bottom with internal anti-vortex equipment |
In quiet conditions, perturbations in the liquid surface are minimal, so gas ingestion is small. However, rolling can be imposed to the stage and transmitted to the drained propellants by the launcher attitude control system,and sloshing can occur. A vortex can form with an associated residual mass that increases with the intensity of these disturbances. To compensate for this,anti-vortex devices have been added to LOX tanks since the development of the Ariane 4 upper stage in order to delay the ingestion of bubbles in the collector and reduce the mass of the diphasic residuals. The impact of these devices is now being computed for a number of different flight configurations with different degrees of rolling associated with different lateral perturbations. These computations are being performed with the complete3D geometry, starting from the very beginning of the launch phase and finishing at the end of the upper stage thrust phase, when the tank is draining.
 Three seconds after the onset of a smooth dip in the surface, a stable vortex is created as the free surface approaches the anti-vortex device. Some bubbles escape from the free surface, and this instant corresponds to the end of the engine feeding process; the remaining liquid in the tank corresponds to the diphasic residual |
 The last phenomenon to be observed is the massive gas ingestion in the collector, which always occurs at the same angular location, due to the curvature of the collector |
The computations focus on two periods of time. The first is dedicated to the thrust of the first stage, and lasts about 10 minutes. The tank is not draining,but the fluids are subjected to external perturbations, particularly spinning.The second is dedicated to the thrust of the second stage, when the tank is draining and subjected to external perturbations. Strong vortex formation appears at the end of this phase with bubble ingestion. This phase lasts about 15 minutes.
Preliminary studies have included a comparison of numerical and experimental results, with tests performed on sub-scale tanks filled with water and at conditions with a similar Froude number. Very good agreement was achieved, from quantitative and qualitative points of view. For example, when the first bubble was ingested, the predicted residual mass of water was within 10% of the measured value. In addition, the qualitative behavior of the free surface in the numerical computations matched the observed behavior, both in terms of the time when dips in the surface were seen to occur and the location and size of the subsequent vortices. Following the validations, further computations were carried out using an actual flight configuration.
Summary
FLUENT is now used intensively to compute the pressure and temperature evolution in cryogenic tanks during all flight phases from pressurization on the launch pad to the last droplet ingestion in the turbo-pump. In the future,the tool will be adopted for use in a micro gravity environment. For this purpose, the development of new local models is needed, and validation of these models will be a difficult task. Non-dimensional numbers will be heavily used to recreate experiments correctly on earth, since cryogenic experiments in true micro gravity conditions are difficult and expensive to perform. Despite the fact that transient 3D computations are CPU intensive,the final goal is to perform them with complex internal geometry coupled with thermodynamic analysis and heat and mass transfer at the wall and at the liquid/gas interface. After complete validation of the relevant models,computations such as these could be performed within a decade.
All these activities have been performed with support from CNES.
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