Ognjan Bozic

Some Aspects Concerning the Design of Multistage Earth Orbit Launchers Using Electromagnetic Acceleration

This paper reports on reflections undertaken to establish a possible multistage space-launch system based on the use of combined electromagnetic and rocket launch technology. It is assumed that the direct access to space is not possible using a single-stage system only, since for such a system, muzzle velocities of approximately 10 km/s are required. Multistage chemical rockets have been successfully used for Low Earth Orbit missions. They are now a mature technology; continued system efficiency improvements are expected to yield only marginal progress. Alternative technology paths, such as the Lorentz rail accelerator (LRA), hold the potential for considerable improvements in first-stage efficiencies compared to rocket systems. Such unconventional solution has serious consequences for the accelerator and payload carrier design. In the first part of this paper, some general aspects (size, forces, energy, and power) of an electromagnetic launcher, which could be able to achieve the required muzzle momentum for such a mission, are discussed. More specifically, the design of such a launcher with respect to the type of electromagnetic LRAs (known also as railguns) and to the armature technology is considered, and also, some of the interstage aspects of such a system are investigated. Special attention is paid to solutions increasing the launch efficiency of the electromagnetic launcher. The second part of this paper considers possible design properties of the rocket carrier (booster and kickoff stage). Propulsion system, structure under high acceleration, thermal protection system (TPS), carrier aerodynamics, and flight mechanic properties are analyzed. The performed analysis confirms that the rocket carrier can be realized with the present technology within the next decade. The conclusion is valid under the assumption that a start velocity from a ground-based electromagnetic LRA is limited to about 4.5 km/s if an adequate TPS is applied. At higher velocities, the extreme heat loads destroy the payload carrier.

Aerothermodynamic Aspects of Railgun-Assisted Launches of Projectiles With Sub- and Low-Earth-Orbit Payloads

A Railgun-concept for the launch of small payloads into low-Earth orbits is currently under investigation. The first major project milestone foresees a 32-MJ railgun system able to launch nonpropelled projectiles with experiments for high-atmospheric research to an altitude of 115 km. This will be followed by a system to launch single-stage rocket-propelled projectiles to put in orbit nanosatellites using a 3.4-GJ railgun with a length of 180 m. One of the main issues to be investigated is the extreme thermal conditions the projectiles surfaces are exposed to, namely high temperatures and high thermal loads. These are caused by friction and turbulent flow properties, when the projectile rushes through the dense Earths atmosphere in less than 30 s. The important questions to be answered for the design of a hypersonic projectile are the determination of the duration, magnitude, and location of the maximum temperatures with respect to the projectiles surface and its corresponding internal structure. For the determination of the complicated physical conditions, a two-step approach is used. In a so-called predesign phase, the HF3T code of the DLR is applied to estimate time-dependent temperature changes on the projectiles surface and/or within the materials structure as a function of the trajectory parameters. Finally, the three-degree-of-freedom trajectory program TRAJECTORY 3D of DLR is used to generate a detailed time-dependent data set, which includes the values for projectile velocity, Mach and Reynolds numbers, and atmospheric properties as a function of altitude. This data set is exchangeable with the TAU code of DLR for the solution of the Navier-Stokes equations, enabling pseudo-unsteady flow solutions along the trajectory. The solutions resulting from both approaches are then compared with emphasis on the surface temperature distributions. This strategy allows the verification of the feasibility of the proposed design solution for the railgun launched hypersonic projectiles

Flight test results of investigation of acceleration effects on a gun launched rocket engine

Nowadays, much research has been done in the field of electromagnetic-driven Lorentz rail accelerators (LRA). To apply this technology to launch a sophisticated payload carrying vehicle, mechanical loads due to high acceleration have to be taken into account. The German Aerospace Center (DLR) is doing research in the field of a hybrid rocket propelled payload carrier which shall be launched from an LRA. The structure of the hybrid rocket engine (HRE) is the most critical component, regarding mechanical stress. To investigate the effects of high acceleration during the launch from an LRA, an experimental setup is created. An 80% scale model of an HRE is mounted on a 155-mm howitzer shell. The experimental setup is then launched from the Panzerhaubitze 2000 with an acceleration of 3300 g. The model is equipped with strain gauge sensors to determine deformation during the acceleration phase. An acceleration sensor is integrated to measure the acceleration during launch. Data of the strain gauge bridges and the accelerometer are sampled by an electronic device mounted on the projectile, and buffered in the internal RAM. The data are transmitted to a ground station during free flight by a telemetry device, which is mounted on the howitzer shell instead of a fuse. The flight path of the projectile is tracked by different radar stations to determine the impact point, so that the experiment can be recovered. The test shows that the HRE structure can withstand the mechanical loads that are caused by high acceleration that would occur during a launch from an LRA. Therefore, an HRE is suitable for high-acceleration launch. The sampled data are compared with finite element method calculations. Differences in simulation and measurement are observed.

Development of an Eddy Dissipation Model for the use in Numerical Hybrid Rocket Engine Combustion Simulation

Within this work, an Eddy Dissipation Model for the combustion process in hybrid rocket engines was developed and implemented into the DLR TAU-Code. The model was developed especially for the propellant combination of hydroxyl-terminated polybutadiene and high concentrated hydrogen peroxide. A numerical generic hybrid rocket combustion chamber was designed to compare the results of the Eddy Dissipation Model simulations with Arrhenius based combustion models. For example, the well validated multistage combustion model from Westbrook and Dryer was applied. The results of the Eddy Dissipation Model simulations are very close to the multistage combustion results and require significantly reduced computing time.

Thermal protection-, aerodynamics- and control simulation of an electromagnetically launched projectile

In recent years several ideas came up to apply electromagnetic launch technology for spaceflight applications. Using electric energy to propel a payload carrier promises the saving of propellant and therefore cost reduction for the transfer to orbit. The conducted studies mostly comprise of a rough estimation of the launcher and the vehicle size. Sometimes a Δv-budget is given to illustrate the energy expenditure. Some of the studies neglect the necessity of a rocket engine. Only by means of an electromagnetic launch without the capability to maneuver, an orbit is not achievable. The high acceleration and the high velocities at low altitude evoke high demands on the payload carrying vehicle. Its structure has to withstand the high acceleration forces during launch and the tremendous aerodynamic heat fluxes during the ascent flight in the dense atmosphere. Moreover a propulsion system, an attitude control system, and a flight controller are needed to bring the vehicle into a circular orbit. This paper presents a vehicle concept that addresses all these demands. The vehicle comprises of a two stage hybrid rocket engine system, a thermal protection system, high test peroxide monopropellant thrusters for attitude control, and a guidance, navigation and control system. A simulation model is created, that consists of a 6-DOF flight mechanics module, aerodynamics model, propulsion module, thermal protection system simulation, as well as of guidance and flight control simulation. Therefore, the complete ascent with all its aspects can be simulated. The simulation results show that a 710 kg vehicle launched with 2586 g and an initial velocity of 3642 m/s can carry 31.5 kg of payload into a 300 km circular orbit. The configuration of the vehicle can be defined by a set of input parameters. This allows using the model within an optimization tool.