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During the start-up of the propulsion system of a satellite or spacecraft, the opening of the tank isolation valve will cause the propellant to flow into an evacuated feedline and slam against a closed thruster valve. This filling process, called priming, can cause severe pressure peaks that could lead to structural failure. In the case of monopropellants such as hydrazine, also, the risk of adiabatic compression detonation must be taken into account in the design of the feedline subsystem. The phenomenon of priming involves complex two-phase flow: the liquid entering the evacuated pipe undergoes flash evaporation creating a vapor cushion in front of the liquid that mixes with the residual inert gas in the line. Moreover, the dissolved pressurizing gas in the liquid will desorb making the priming process difficult to model. In order to study this phenomenon, a new test-bench has been built at DLR Lampoldshausen which allows fluid transient experiments in the same conditions as the operating space system. Tests are performed with water and ethanol at different conditions (tank pressure, vacuum level, pressurizing gas helium vs. nitrogen, etc.). The effect of the geometry is also investigated, comparing different test-elements such as straight, tees, and elbow pipes. The pressure profile is found to be dependent on the geometry and on the downstream conditions. The acoustic wave reflection caused by the pipe geometry and fluid dynamic effects such as the aforementioned desorption and flash evaporation induce a complex pressure profile of the first pressure peak. Finally, numerical simulations of the priming process are performed by means of EcosimPro software in conjunction with European Space Propulsion System Simulation (ESPSS) libraries and results are compared with experiments.

Fluid hammer occurs when a flow in a pipeline is rapidly stopped due to valve closure or shutting down a pump. The resulting pressure wave can reach very high amplitudes, depending on the fluid properties and the flow velocity. In spacecrafts and launchers, where the reduction of weight is always an important design goal, it is not possible to build the structure arbitrarily robust. Understanding the transient behaviour of the fluids is necessary to predict mechanical loads on the structure and align the design to them. Since a lot of launchers use reactive, cryogenic propellants, it is a common approach to use inert substitute fluids for on ground testing like water (H2O) or liquid nitrogen (LN2). LN2 comes with the advantage of being cryogenic like real propellants, but ground testing is consequently more complex than with H2O. For this purpose, several fluid hammer experiments with both fluids were performed and compared to each other to provide a foundation for deciding which substitute fluid would be useful. Graphical abstract

A thermal lattice-Boltzmann model of a van der Waals gas was used to check its applicability to specific challenges in the numerical simulation of transient cryogenic two-phase flow in rocket engine feed systems. Three test cases were chosen to prove the model is capable of capturing the underlying physics. Overall correct representation of incompressible flow should be demonstrated by a lid-driven cavity. The capability of the model to handle shocks and supersonic flow is shown in a shock tube configuration. The last test case was chosen in order to show instantaneous evaporation by the formation of a single vapour bubble at a heated surface.

The DLR LUMEN (Liquid Upper stage deMonstrator ENgine) rocket engine comprises of two separate turbopumps, one each for Liquid Oxygen (LOX) and Liquified Natural Gas (LNG) supply. In both turbopumps, an identical bearing block separates the pump section from the turbine section. In the current design, the bearing block features an oil-jet lubrication system to cool and lubricate the bearings during operation. Cryogenic conditions at the pump-interface and high temperature conditions at turbine-interface impose thermal constraints on the selection of a suitable lubricant for the bearing assembly and the design of the lubrication system. As a starting point to address this challenge, this paper presents a preliminary uncoupled finite element thermal model of the LUMEN LOX turbopump. The thermal model investigates the thermal behavior of the housing and rotor components in the bearing housing from start-up to shutdown of the turbopump. The model employs thermal worst case conditions to establish the operational thermal envelope of the turbopump system. Spatial and temporal evolution of the temperature at critical points within the bearing housing are reported. The model, albeit the uncertainties rooted from its empirical nature, provides beneficial insight into the thermal characteristics of the turbopump assembly. This serves as a first estimate of the thermal constraints for the selection of lubricant oil.

In rocket engines seals are one of the key elements for a successful turbopump operation. They are not only responsible for reducing the leakage but they also have a huge influence on the rotordynamic behaviour. Depending on the design, the seal on a turbopump impeller can damp or excite rotordynamic instabilities. In this work we present transient 3D CFD simulations of straight gap seals, which are used to determine the rotordynamic coefficients. After that the numerical simulation is extended to characterize the influence of the impeller to housing clearance and swirl brakes on the rotor dynamic coefficients. We show that the inertia term cannot be neglected for the investigated nominal rotational speed of the rotor and high absolute whirl frequency ratios especially for the rotor housing clearance. Finally the coefficients calculated by numerical simulation are used in the equations of motion for a floating ring seal and the equations are solved numerically to do a sensitivity study of a floating ring to variations in axial force, rotor displacement and floating ring mass.