VINCI®, the european reference for ariane 6 upper stage cryogenic propulsive system

The intent of this publication is to provide an overview of the development of the VINCI® engine over the period 2014–2015. The VINCI® engine is an upper stage, cryogenic expander cycle engine. It combines the required features of this cycle, i. e., high performance chamber cooling and high performance hydrogen turbopump, with proven design concepts based on the accumulated experience from previous European cryogenic engines such as the HM7 and the VULCAIN®. In addition, its high performance and reliability, its restart and throttle capability offer potential applications on various future launcher upper stages as well as orbital spacecraft. At the end of 2014, the VINCI® successfully passed the Critical Design Review that was held after the major subsystem (combustion chamber, fuel and oxygen turbopump) had passed their own Critical Design Review all along the second half of 2014. In December, a Ministerial Conference at government level gave priority to the Ariane 6 program as Europe future launcher. In the framework of this decision, VINCI® was confirmed as the engine to equip Ariane 6 cryogenic upper stage engine. This publication shows how the VINCI development is progressing toward qualification, and also how the requirements of the new Ariane 6 configuration taken into account, i. e., offering new opportunities to the launch system and managing the new constraints. Moreover, the authors capitalize on the development already achieved for the evolution of Ariane 5. In parallel to completing the engine development and qualification, the configuration and the equipment of the propulsive system for Ariane 6 such as the components of the pressurization and helium command systems, board to ground coupling equipment, are being defined.

Quasi-two-dimensional approximation for numerical simulation of flows in engine ducts

New quasi-two-dimensional (2.5D) approach to description of three-dimensional (3D) flows in ducts is proposed. It generalizes quasi-one-dimensional (quasi-1D, 1.5D) theories. Calculations are performed in the (x; y) plane, but variable width of duct in the z direction is taken into account. Derivation of 2.5D approximation equations is given. Tests for verification of 2.5D calculations are proposed. Parametrical 2.5D calculations of flow with hydrogen combustion in an elliptical combustor of a high-speed aircraft, investigated within HEXAFLY-INT international project, are described. Optimal scheme of fuel injection is found and explained. For one regime, 2.5D and 3D calculations are compared. The new approach is recommended for use during preliminary design of combustion chambers.

A preliminary design study for an expander liquid oxygen turbopump

In the recent years, Airbus DS GmbH started a turbopump initiative to buildup fundamental capabilities in analyzing and designing turbomachinery within a German national funded program “TARES.” Turbomachinery is widely used in different rocket propulsion systems and include such parts as pumps and turbines. Turbines are used for generating power required by pumps in order to feed the propellants to the thrust chamber. The paper is dedicated to present an overview about currently ongoing conceptual design activities of turbomachinery covering the main design phases like TPA (TurboPump Assembly) layout tradeoff; rotational speed selection with respect to efficiency and cavitation; flow path design techniques including blade profiling; computer-aided design (CAD) work; and preliminary structural analyses. This paper presents the main outcome applying the established design logic to a liquid oxygen (LOx) turbomachinery. The component is designed based on a dedicated specification for an expander cycle type engine. This includes a LOx pump unit comprising inducer and impeller as well as a subsonic single stage reaction turbine. For the turbine drive, gaseous hydrogen (GH2) heated within the thrust chamber cooling circuit is used. Within this paper, a general overview about the preliminary work results of pump and turbine sizing, profiling, performance estimation as well as structural aspects is given.

Flame response to high-frequency oscillations in a cryogenic oxygen/hydrogen rocket combustor

Experiments presented in this paper were conducted with the BKH rocket combustor at the European Research and Technology Test Facility P8, located at DLR Lampoldshausen. This combustor is dedicated to study the effects of high magnitude instabilities on oxygen/hydrogen flames, created by forcing high-frequency (HF) acoustic resonance of the combustion chamber. This work addresses the need for highly temporally and spatially resolved visualization data, in operating conditions representative of real rocket engines, to better understand the flame response to high amplitude acoustic oscillations. By combining ONERA and DLR materials and techniques, the optical setup of this experiment has been improved to enhance the existing database with more highly resolved OH* imaging to allow detailed response analysis of the flame. OH* imaging is complemented with simultaneous visible imaging and compared to each other here for their ability to capture flame dynamics.

Large-eddy simulation of transcritical liquid oxygen/methane jet flames

A numerical method to perform large-eddy simulations (LES) of nonpremixed liquid oxygen/methane (LOx/CH4) combustion at supercritical pressures is presented and the computational results are compared with available experimental data. The injection conditions of the considered test case resemble those in typical liquid-propellant rocket engines (LRE). Thermodynamic nonidealities are modeled using the Peng–Robinson (PR) equation of state (EoS) in conjunction with a novel volume-translation method to correct deficiencies in the transcritical regime. The resulting formulation is more accurate than the standard cubic EoS's without deteriorating their good computational efficiency. The real-gas thermodynamics model is coupled with the steady laminar flamelet model (SLFM) for turbulent nonpremixed combustion to incorporate chemical reactions at reasonable computational cost in the LES. A reduced reaction mechanism, which is validated with respect to the full mechanism, is used to generate a flamelet library. A comparison of the LES result with available OH* measurements shows that important flow features are well predicted.

One-dimensional model describing eigenmode frequency shift during transverse excitation

A shift in transverse eigenmode frequency was observed in an experimental combustion chamber when exposed to large amplitude acoustic oscillations during oxygen–hydrogen combustion tests. A shift in eigenmode frequency under acoustic conditions representative of combustion conditions is of critical importance when tuning acoustic absorbers or investigating injection coupled combustion instabilities. The experimentally observed frequency shift was observed both in the frequency domain and as an asymmetric amplitude response to a linear frequency ramp of an external excitation system in the time domain. The frequency shift was found to be dependent on amplitude and operating condition. A hypothesis is presented for the frequency shift based on change in speed-of-sound distributions due to flame contraction when exposed to high amplitude pressure oscillations. A one-dimensional (1D) model was created to test the hypothesis. Model parameters were based on relationships observed in experimental data. The model was found to accurately recreate the frequency shifting asymmetric response observed in test data as well as its amplitude dependence. Further development is required to investigate the influence of operating conditions and chamber design on the quantitative modeling of the frequency shift.

Green gelled propellant highly throtteable rocket motor and gas generator technology: status and application

Work on gel propulsion began in Germany in 1999. The German Gel Propulsion Technology (GGPT) Program started in 2001 from a white sheet of paper, proposed by the DLR Institute of Space Propulsion, Bayern-Chemie (BC), and the Fraunhofer Institute of Chemical Technology. Aim of the first phase was to develop the technology needed to build a rocket motor burning gelled propellants and to demonstrate its operability by a free flight within significantly less than a decade. The research and development activities were guided by a suitable principal concept for a gelled propellant rocket motor (GRM). Based on theoretical considerations (regarding functional aspects) and experimental pre-tests (propellant development, gelation, rheology, spraying, ignition, and combustion), a motor concept was pre-selected and the motor developed. The identified requirements were proven in December 2009 by two successful demonstration flights. The achieved know-how from basic research and technology (R+T) development has been consolidated in an application-oriented way on component level up to motor development. Within this scope, also, the goal of an effective control of the thrust by throttling the propellant mass flow rate (PMFR) while maintaining an optimum combustion chamber (CC) pressure could be achieved. This publication describes briefly several major advances in the development of the gel propulsion technology in Germany from rheology to combustor development to the thrust and pressure controller of a GRM.

Hot-fire testing of liquid oxygen/hydrogen single coaxial injector at high-pressure conditions with optical diagnostics

Injector behavior is of utmost importance for the performance and stability of liquid rocket engines (LREs). A major problem is getting a highly efficient homogeneous mixture and effective chemical reaction of fuels at minimum chamber length. Despite substantial progress in numerical simulations, a need for experimental data at representative conditions for development and validation of numerical design tools still exists. Therefore, in the framework of the DLR-project “ProTau,” the authors have performed tests to create an extended data base for numerical tool validation for high-pressure liquid oxygen (LOx) / hydrogen combustion. During the experimental investigations, a windowed DLR subscale thrust chamber model “C” (designated BKC) has been operated over a broad range of conditions at reduced pressures of approximately 0.8 (4 MPa), 1 (5 MPa), and 1.2 (6 MPa) with respect to the thermodynamic critical pressure of oxygen. Liquid oxygen and gaseous hydrogen (GH2) have been injected through a single coaxial injector element at temperatures of ~ 120 and ~ 130 K, respectively. High-speed optical diagnostics have been implemented, including imaging of OH* emission and shadowgraph imaging at frequencies from 8 up to 10 kHz to visualize the flow field.

Experimental study of water hammer pressure surge

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.

Analytical viscous flow model and test validation of a swirl injector for rocket engine application

The generally adopted flow model inside a swirl injector, widely used injection concept for propulsive applications, relies upon the hypothesis of ideal flow neglecting the fluid viscosity effects. This model showed significant prediction errors with relatively high viscosity propellants, often leading to the need of an experimental characterization of the injection elements. In this paper, an analytical approach is presented, which includes the effects of viscous diffusion on the injector performance leading to a close form flow solution. The built model is thus experimentally validated testing a liquid oxygen (LOx) and an ethanol injector: the good agreement between the model and the experimental results leads to the construction of the injectors operational maps describing the injector behavior even in the presence of viscous effects.