Dario Giuseppe Pastrone

Optimal Design and Control of Hybrid Rockets for Access to Space

A nested direct/indirect method is used to find the optimal design of hybrid rockets for suborbital missions. The direct optimization of the parameters which affect the engine design is coupled with the trajectory indirect optimization. Different propellant combinations are analyzed. First, the simplest blowdown feed system is considered. A more complex pressurization system with an additional gas tank, which allows a phase with constant propellant tank pressure, is then analyzed. The optimization procedure provides the engine design and the trajectory, which maximize the mission performance index.

Optimal Design of Hybrid Rocket Motors for Launchers Upper Stages

A hybrid rocket is considered as the third stage of a three-stage launcher. The propulsion system design and the trajectory are simultaneously optimized by means of a nested direct/indirect procedure. Direct optimization of the parameters that affect the motor design is coupled with indirect trajectory optimization to maximize the launcher payload for assigned conditions at the stage ignition and final orbit. A mission profile based on the Vega launcher is considered. The feed system exploits a pressurizing gas, namely helium, with hydrogen peroxide as the oxidizer and polyethylene as the fuel. The simplest blowdown design is compared with a more complex pressurizing system, which has an additional gas tank that allows for a phase with constant oxidizer tank pressure. The optimization provides the optimal values of the main engine design parameters (pressurizing gas mass, nozzle expansion ratio, and initial values of tank pressure, mixture ratio and thrust), the corresponding grain and engine geometry, and the control law (thrust direction during the ascent trajectory and engine switching times). Results show that a hybrid rocket may be a viable option for small launchers.

Design Trade-offs for Hybrid Rocket Motors

A new method is developed for preliminary design of hybrid rocket motors. A cooperative evolutionary method is used with an indirect approach to perform the coupled optimization of hybrid rocket motor and trajectory for a sounding rocket. Different propellant combinations are considered for a sounding rocket application (microgravity platform). The motor grain is cylindrical with a single circular port: blowdown and self-pressurized feeding systems are used. The eects of throat erosion are discussed, taking into account recent CFD evaluation of throat regression rate in nozzles of hybrid rocket motors. The method shows to be very promising with respect to other procedures and can be applied to complex applications such as launcher upper stages or landers. Due to the peculiarity of the present problem also a purely evolutionary approach is used to perform optimization. Even if the nested evolutionary/indirect method is able to find a better solution, the purely evolutionary method is able to find a near-optimal solution with a CPU time which is one order of magnitude lower

Optimization of a Hybrid Rocket Upper Stage with Electric Pump Feed System

An electrical pump is considered to feed the oxidizer into the combustion chamber of a hybrid rocket motor, which is used as the third stage of a three-stage launcher. The motor uses hydrogen peroxide as the oxidizer and polyethylene as the fuel; advanced Lithium batteries are adopted to power the pump. The design of the hybrid rocket motor and the trajectory are simultaneously optimized by means of a nested direct/indirect procedure. Direct optimization of the parameters that affect the motor design is coupled with indirect trajectory optimization to maximize the launcher payload for assigned characteristics of the first and second solid propellant stages and final orbit. The optimization provides the optimal values of the main engine design parameters, the corresponding grain and engine geometry, and the control law. A mission profile based on the Vega launcher is considered. The performance obtained using an electrical pump feed system is compared with pressuregas feed systems. Results show the relevant improvements that can be obtained when a present-technology electrical pump feed system is adopted.

Hybrid Evolutionary Method for Optimization of Mixed-Flow Civil Turbofan

A mixed-stream turbofan is considered for a civil transport; an optimization procedure based on a hybrid evolutionary algorithm, which employs a genetic algorithm, differential evolution and particle swarm optimization in parallel, is used to minimize the aircraft take-off gross weight, given payload, take-off and landing distance, time to climb to cruise altitude, and range. The optimization procedure provides the optimal values of wing loading and engine design variables; in particular, the optimal mixing condition is determined. The effect of the engine weight estimation on the results is investigated. Results confirm that the benefit of mixing core and bypass streams can be appreciated when moderate values of the bypass ratio are allowed, whereas it may become marginal for large values of the bypass ratio.

Integrated Design of Hybrid Rocket Upper Stage and Launcher Trajectory

A three-stage launcher, with solid-propellant first and second stage and a hybridpropellant third stage is considered. The design of the hybrid-propellant upper stage and the whole ascent trajectory are simultaneously optimized by means of a nested direct/indirect procedure, while the characteristics of the first and second stages are assigned. Direct optimization of the parameters that affect the motor design is coupled with indirect trajectory optimization to maximize the launcher payload. A mission profile based on the Vega launcher is considered. The feed system exploits a pressurizing gas, namely helium, with hydrogen peroxide as the oxidizer, and polyethylene as the fuel. The simplest blowdown design is compared with a more complex pressurizing system, which has an additional gas tank that allows for a phase with constant pressure in the oxidizer tank. The optimization provides the optimal values of the main engine design parameters (pressurizing gas mass, nozzle expansion ratio, and initial values of tank pressure, mixture ratio and thrust), the corresponding grain and engine geometry, and the control law (thrust direction during the ascent trajectory and engine ignition and shutoff times). Results show that a hybrid-propellant third stage may be a viable option for small launchers, with improved performance and similar cost compared to an all-solid rocket. The results of the optimization also offer interesting theoretical insight into the problem.

Optimal Design of Hybrid Rockets with Self-Pressurizing Oxidizer

A nested direct/indirect method is used to find the optimal design of a microgravity platform which consists of a hybrid sounding rocket that uses a self-pressurizing oxidizer, namely, nitrous oxide. The direct optimization of the parameters that affect the engine design is coupled with the trajectory indirect optimization to maximize a given mission performance index. Different models can be used to describe the self-pressurizing behavior of the oxidizer in the tank. The simplest model assumes liquid/vapor equilibrium. A two-phase model is also proposed: saturated vapor and superheated liquid are considered and the liquid/vapor mass transfer is evaluated making reference to the liquid spinodal line. Results show that the different models have a limited impact in the optimal engine characteristics. The performance are slightly modified due to the different mass of the residual oxidizer. A performance comparison with different propellant combinations is also shown.

Optimization of Hybrid Propellant Mars Ascent Vehicle

The use of hybrid rocket engines for ascent from Mars’s surface is analyzed. The engine design and the ascent trajectory are optimized by means of a nested procedure which employs a direct method to optimize the engine design parameters and an indirect method to optimize the ascent trajectory. Ascent trajectories of a two-stage vehicle for the return of samples from Mars’ surface and for a manned mission are considered as test cases. The same engine is used in both the first stage (a cluster of engines is considered) and for the second stage (a single engine) in order to limit the development costs. Results demonstrate the feasibility of the concept.

Optimization of Civil Turbofan with Evolutionary Algorithms

A high-bypass ratio turbofan is considered for a civil transport; an optimization procedure based on a hybrid evolutionary algorithm, which employs a genetic algorithm, differential evolution and particle swarm optimization in parallel, is used to minimize the aircraft take-off gross weight, given payload, take-off and landing distances, time to climb to cruise altitude, and range. Different performance indexes are also considered for the sake of comparison. The procedure provides the optimal values of wing loading and engine design variables. Off-design analysis is embedded into the optimization code to asses the fulfillment of performance constraints.

Optimal mixture-ratio control for a single-stage-to-orbit rocket

An indirect procedure is used to optimize the performance of a Single-Stage-To-Orbit rocket which uses liquid oxygen and hydrogen as propellants. A simplified model for the ascent trajectory is considered to highlight the effects of the propellant mixture-ratio control. By assuming that the tank mass is proportional to the propellant volume, either the gross mass or the dry mass are minimized for an assigned payload. The results quantify the advantage that can be obtained by using a high mixture ratio (i.e., high propellant bulk density) in the initial phase of the trajectory and by exploiting a low mixture ratio (i.e., high specific impulse) in the final phase. Because of the ascent losses, the control law is different to a deep space operation, where gravitational and aerodynamic forces are not present. A higher mixture ratio is used in the initial phase to obtain a slightly higher thrust level and a faster vehicle mass reduction. The continous mixture-ratio control provides a slight benefit in terms of propellant consumption compared to the constant mixtureratio operation, while a larger benefit is obtained in terms of payload, due to the tank mass reduction.