Gerald Hagemann

ADVANCED ROCKET NOZZLES

Several nozzle concepts that promise a gain in performance over existing conventional nozzles are discussed in this paper. It is shown that signie cant performance gains result from the adaptation of the exhaust e ow to the ambient pressure. Special attention is then given to altitude-adaptive nozzle concepts, which have recently received new interest in the space industry. Current research results are presented for dual-bell nozzles and other nozzles with devices for forced e ow separation and for plug nozzles with external freestream expansion. In addition, results of former research on nozzles of dual-mode engines such as dual-throat and dual-expander engines and on expansion ‐dee ection nozzles are shown. In general, e ow adaptation induces shocks and expansion waves, which result in exit proe les that are quite different from idealized one-dimensional assumptions. Flow phenomena observed in experiments and numerical simulations during different nozzle operations are highlighted, critical design aspects and operation conditions are discussed, and performance characteristics of selected nozzles are presented. The consideration of derived performance characteristics in launcher and trajectory optimization calculations reveal signie cant payload gains at least for some of these advanced nozzle concepts.

Restricted Shock Separation in Rocket Nozzles

In overexpanded rocket nozzles the e ow separates from the nozzle wall at a certain pressure ratio of wall pressure to ambient pressure. Flow separation and its theoretical prediction have been the subject of several experimental and theoretical studies in the past decades. Two distinctive e ow separation phenomena, the freeshock and restricted-shock separation, were observed in experiments with nozzles. Both phenomena are discussed in detail, and the system of recompression shocks and expansion waves is described. For the free-shock case three different shock structures in theplume can occur, namely the regular shock ree ection, the Mach disk, or a cap-like shock pattern. Theappearanceofthesedifferentplumepatternsis discussed. Theseshock structuresareconserved for the full-e owing, but overexpanded, nozzle. Numerical results obtained for existing rocket nozzles, e.g., Space ShuttleMain EngineorVulcain, show a qualitativegood agreement with experimental photographs.Furthermore, an explanation for the appearance of restricted shock separation, which has been widely unknown up to now, is given, analyzing why and under what conditions it occurs. The type of nozzle contour strongly ine uences this form of e ow separation, and restricted shock separation also occursin full-scale, thrust-optimized rocket nozzles. Based on the results established for e ow separation, an outlook on the generation of side loads is given.

Critical Assessment of Dual-Bell Nozzles

A critical assessment of dual-bell nozzles is given in this paper. The principal flow field development in dual-bell nozzles, as well as design aspects for the contour of the base nozzle, the wall inflection, and the nozzle extension are discussed. Special regard is focused on the transition behavior from sea level to altitude operation and its dependence on the contour type used for the nozzle extension. Parametric numerical simulations of the flowfield development were performed to quantify the different loss effects. It is shown that the additional performance losses caused by the dual-bell nozzle contour are surprisingly low. An analytical derivation of the flow transients from the separated to the fully attached flow is presented. The necessity of further experimental investigations on dual-bell nozzles is emphasized, which will lead to a better understanding of the flow transition in dual-bell nozzles. Finally, new ideas are presented to minimize the duration of the critical flow transition by varying the thrust chamber pressure on system level, to ensure a sudden and controlled jump of the separation point from the wall inflection (sea-level operation) to the exit plane (altitude operation).

Critical Assessment of the Linear Plug Nozzle Concept

The linear plug nozzle concept has been investigated by means of analytical and experimental work. Different design approaches for annular and linear plug applications published in the literature are critically discussed. Based on earlier analytical and experimental work published by the authors, a new design methodology for linear plug nozzles with combined internal and external expansion was developed and experimentally validated. This experimental work included cold gas subscale tests performed within the frame of the German National Technology Programme LION. Complementary tests were performed within a Russian National Programme to quantify additional losses due to clustering. It is shown that depending on the contour type used, a smooth altitude mode operation can be achieved. In addition, the performance of a round-to-square nozzle used as primary expansion module is investigated.

Scalability for Rocket Nozzle Flows Based on Subscale and Full-Scale Testing

When flow separation and side-load behavior is investigated for different nozzle concepts or when design considerations for advanced nozzle concepts are experimentally screened, the initial experimental effort is generally limited to subscale experiments with high-quality diagnostics. However, one fundamental question always arises: How should these test results be scaled to a potential full-scale design or application? Different approaches exist, including the use of characteristic numbers for normalization and analytical and numerical models. It is shown that analytical models, being based on physical origins and dominated by inviscid flow phenomena, may result in high accuracy for full-scale prediction. In contrast, it is shown that for flow regions dominated by viscous effects significant differences between cold- and hot-gas experiments are observed, which makes a direct scaling from cold-gas subscale to hot-gas full-scale difficult. For scaling, an intermediate step in the scaling rationale with hot-firing tests is foreseen. Demonstrated by results from hot-firing tests with Vulcain subscale nozzles, it is shown that a geometrical scaling is possible for similar hot-gas flow properties. These test results, obtained with a 40-kN thrust chamber at representative combustion chamber operation conditions, are presented and discussed in detail.

LOX-Hydrocarbon Preparatory Thrust Chamber Technology Activities in Germany

Recent European interest in LOX/hydrocarbon propulsion systems focuses on LOX/kerosene versus LOX/methane propellants for liquid booster stages. For the preparation of a fair technical trade-off among the two propellant combinations, EADS-ST and DLR have started complementary activities addressing critical areas for the related thrust chamber development. These activities are closely coordinated, and are targeting on fundamentals such as the characterization of the propellant fluid properties, the propellant cooling characteristics, material characteristics, and ignition studies. Besides these fundamental research activities, further technology activities focusing on future application towards open and closed cycle engines have been started. This includes experimental verification of injector head designs, but also the enhancement of design tools to predict the hot gas side and coolant side heat transfer at relevant operational conditions. The technology activities finally focus on staged combustion battleship demonstrators for the coupled thrust chamber / preburner system, allowing for a fair technical trade-off between both hydrocarbon fuels.

LOX/Hydrocarbon Propellant Trade Considerations for Future Reusable Liquid Booster Engines

Recent investigations in Europe on future space transportation have traded among others the pro’s and con’s of LOX/kerosene versus LOX/methane propellants for liquid booster stages, including both vehicle and engine aspects. On vehicle side, both hydrocarbon propellants were compared for reusable first stages and reusable as well as expendable ARIANE 5 liquid booster substitutes. On engine side, the propellant trade was focusing on kerosene and methane characteristics in regard to performance and thermodynamic cycles, key subsystem pre-design and life potential, as well as in regard to the maturity status of the engine technology to be employed. The results highlighted that on vehicle level, no crucial advantage especially in view of payload performance could be identified for either kerosene or methane. On engine level, methane indicated to be superior to kerosene in regard to subsystem life potential in particular w.r.t. the main combustion chamber, but clearly lacked an adequate technology maturity level as demonstrated by the kerosene engines being in operation today.

Optimization of dual-expander rocket engines in single-stage-to-orbit vehicles☆

Abstract Dual-expander rocket engines offer a trajectory adapted dual-mode operation during the ascent of a launcher, which may be of significant advantage for single-stage earth-to-orbit vehicles, when compared to conventional rocket engines with bell-type nozzles. This paper investigates a reusable single-stage earth-to-orbit vehicle with a constant payload capability of 16.5 Mg into low earth orbit, for the comparison of the dual-expander rocket engines with conventional rocket engines, using only hydrogen and oxygen as propellant combination in all engines.

TEKAN 2010 - Thrust Chamber Technologies for Liquid Rocket Propulsion

Abstract Facing the worldwide competition in the space transportation market, Europe today concentrates its efforts on cost reduction for future production batches of its Ariane launcher family safeguarding its leading position as commercial space access provider. In addition to this, alternative launcher scenarios are evaluated in line with associated key enabling technologies. These investigations include complementary expendable launch systems as well as long-term reusable launch systems, the latter being primarily considered within the European Space Agency’s Future Launcher Preparatory Programme FLPP. Within TEKAN 2010, alternative two stage ELV scenarios are being analysed with a special focus on 1 st -stage propulsion systems comprising both gas-generator and staged-combustion engine options. Enabling technologies for the thrust chamber assembly system are addressed, offering dual-use in terms of cost reduction and cycle application. Status of injector technology, thrust chamber manufacturing technology and further component technology is discussed aiming at a TRL of 4 to 5 at project completion in 2007. Furthermore, the evolution of advanced engineering design tools is briefly highlighted.

Main Stage Liquid Propulsion Acitivities within Europe's Future Launcher Preparatory Programme FLPP

In 2004, the ESA Council decided to start the Future Launcher Preparatory Programme, FLPP. Objective of this programme is to prepare Europe for the decision on its Next Generation Launcher within the next years to come. Within the current programme status, both reusable and expendable concepts are studied. Facing Europe’s heritage on expendable launch vehicles, it has been decided to concentrate the initial effort towards reusable concepts and on key-enabling technologies meeting ambitious vehicle and its subsystem service life requirements. A requirement driven development approach has been adopted to define the propulsion system requirements and to derive the further need for technology demonstration by ground tests. Three reference engine concepts have been worked out by the propulsion consortium, one for each fuel combination Hydrogen, Methane and Kerosene. Technology work started on identified critical subsystems, including turbopumps, valves, staged combustion demonstration, and HMS. Details are presented in this paper.