Robust Design Approaches for Hybrid Rocket Upper Stage

Computational costs of robust-based design optimization methods may be very high. Evaluation of new procedures for the management of uncertainty with applications to hybrid rocket engines is here carried out. Two newly developed procedures are presented (hybrid algorithm and iterated local search), and their performances are compared with those of two previously developed procedures (genetic algorithm and particle swarm optimization). A liquid oxygen/paraffin-based fuel hybrid rocket engine that powers the third stage of a Vega-like launcher is considered. The conditions at third-stage ignition are assigned, and a proper set of parameters are used to define the engine design and compute the payload mass. Uncertainties in the regression rate are taken into account. An indirect trajectory optimization approach is used to determine a mission-specific objective function, which takes into account both the payload mass and ability of the rocket to reach the required final orbit despite uncertainties. Results show that for this kind of problem, particle swarm optimization and iterated local search outperform the genetic algorithm, but the use of a local search operator may slightly improve its performance

Multi physics modelling for a hybrid rocket engine with liquefying fuel: a sensitivity analysis on combustion instability

Hybrid rocket engines represent a promising alternative to both solid rocket motors and liquid rocket engines. They have throttling and restart capabilities with performance similar to storable liquids, but are safer and are low-cost. However, some drawbacks, such as low regression rate and combustion instability, are limiting their effective application. Paraffin-based fuels are a solution envisaged to face the low regression rate issue, and the capability to describe and predict combustion instability in the presence of liquefying fuels becomes an enabling step towards the application of hybrid rockets in next-generation space launchers. In this work, a multi physics model for hybrid rocket engines is presented and discussed. The model is based on a network of submodels, in which the chamber gas dynamics is described by a quasi-1D Euler model for reacting flows while thermal diffusion in the grain is described by the 1D heat equation in the radial direction. The need to introduce strong modelling simplifications introduces a significant uncertainty in the predictive capability of the numerical simulation. For this reason, a sensitivity analysis is performed in order to identify the key parameters which have the largest influence on combustion instability. Results are presented on a test case which refers to a paraffin-based grain burnt with hydrogen peroxide

Optimal battery selection for hybrid rocket engine

In the present paper, the optimal selection of batteries for an electric pump -fed hybrid rocket engine is analyzed. A two -stage Mars Ascent Vehicle, suitable for the Mars Sample Return Mission, is considered as test case. A single engine is employed in the second stage, whereas the first stage uses a cluster of two engines. The initial mass of the launcher is equal to 500 kg and the same hybrid rocket engine is considered for both stages. Ragone plot -based correlations are embedded in the optimization process in order to chose the optimal values of specific energy and specific power, which minimize the battery mass ad hoc for the optimized engine design and ascent trajectory. Results show that a payload close to 100 kg is achievable considering the current commercial battery technology

Hybrid rocket engine design optimization at politecnico di torino: A review

Optimization of Hybrid Rocket Engines at Politecnico di Torino began in the 1990s. A comprehensive review of the related research activities carried out in the last three decades is here presented. After a brief introduction that retraces driving motivations and the most significant steps of the research path, the more relevant aspects of analysis, modeling and achieved results are illustrated. First, criteria for the propulsion system preliminary design choices (namely the propellant combination, the feed system and the grain design) are summarized and the engine modeling is presented. Then, the authors describe the in-house tools that have been developed and used for coupled trajectory and propulsion system design optimization. Both deterministic and robust-based approaches are presented. The applications that the authors analyzed over the years, starting from simpler hybrid powered sounding rocket to more complex multi-stage launchers, are then presented. Finally, authors’ conclusive remarks on the work done and their future perspective in the context of the optimization of hybrid rocket propulsion systems are reported

Optimal Design of Hybrid Rocket Small Satellite Launchers: Ground Versus Airborne Launch

A three-stage hybrid rocket is considered as a small satellite launcher. The same engine is used in different numbers in each stage: six, three, and one in the first, second, and third stages, respectively. This design choice aims at an overall reduction of the launcher cost. The propellants are liquid oxygen and a paraffin-based fuel. The performance of different feed systems and launch options are evaluated: the feasibility of a ground launch is analyzed and compared to similar three-stage launchers with airborne launch using both a gas pressurized feed system and an electric turbopump feed system. The optimization procedure exploits a direct method to evaluate the best values of engine design parameters, whereas an indirect method optimizes the ascent trajectory once the engine design is given. Constant power and blowdown operation are, respectively, assumed for the electrical feed system and the gas pressurized feed system. The initial mass of the launcher is given (5000 kg), and the payload mass is maximized for a given insertion orbit. The initial thrust is fixed in order to have an initial acceleration equal to 1.4g. The nozzle expansion ratio in the first-stage engines is reduced to avoid separation at liftoff in the ground case, and the third-stage engines are used at a lower vacuum thrust level to satisfy maximum acceleration constraints. The results show that the proposed small satellite launcher concepts are able to deliver payload masses in the range of 40-100 kg into the desired orbit