This paper presents a novel approach to the solution of multiphase multi-objective optimal control problems. The proposed solution strategy is based on the transcription of the optimal control prob...

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Direct Solution of Multi-Objective Optimal Control Problems Applied to Spaceplane Mission Design

A novel, two-way coupled, dusty-gas flow model has been developed in the direct simulation Monte Carlo (DSMC) framework and employed for the dust-dispersion study on lunar surface. In this model, the gas–gas collisions are modeled probabilistically, whereas, grain–grain interactions are computed deterministically. Most importantly, the gas–grain interactions are modeled in a two-way coupled manner through the consideration of momentum and energy exchange between the two phases. The proposed model is validated against the two-phase theoretical relations for a zero-dimensional simulation. The computational model is used to study the dust dispersion problem due to plume impingement on lunar surface. The influence of particle diameter and hovering altitudes on gas and grain phases, and dust transportation are analyzed in the modified DSMC framework. Furthermore, the sensitivity of the two-way coupled gas–grain interaction model is discussed in relation to the one-way coupled model.

Modeling of dusty gas flows due to plume impingement on a lunar surface

Space penetrators are a potential method of inserting instrumentation onto ice-covered bodies in the solar system. Part of a study to see whether this is feasible involves numerically simulating impact of the penetrator into ice at impact velocities of a few 100 m/s. In order to do this accurately, it is necessary to have a constitutive model for water ice that is valid at the strain rates and temperatures relevant to impact in the Outer Solar System. This paper reports certain issues and difficulties that arose during a study to obtain this data.

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Problems Associated with Making Mechanical Measurements on Water–Ice at Quasistatic and Dynamic Strain Rates

A study of the compressive mechanical properties of defect-free, porous and sintered water-ice at low and high strain rates

Conceptual design analysis for a two-stage-to-orbit semi-reusable launch system for small satellites

This paper presents the conceptual design and performance analysis of a partially reusable space launch vehicle for small payloads. The system uses a multi-stage vehicle with rocket engines, with a reusable first stage capable of glided or powered flight, and expendable upper stage(s) to inject a 500 kg payload in different low Earth orbits. The space access vehicle is designed to be air-launched from a modified aircraft carrier. The aim of the system design is to develop a commercially viable launch system for near-term operation, thus emphasis is placed on the efficient use of high TRL technologies. The vehicle design are analysed using a multi-disciplinary design optimisation approach to evaluate the performance, operational capabilities and design trade-offs.

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Vehicle and mission design of a future small payload launcher

Soft landings on extra-terrestrial airless bodies will be required for future sample return missions, such as the Phobos Sample Return (PhSR). PhSR is a candidate mission of ESA’s Mars Robotic Exploration Preparation (MREP-2) Programme. Its main objective is to acquire and return a sample from the Martian moon Phobos, after a scientific characterisation phase of the moon and of the landing site. If a rocket is used to slow down the spacecraft to a vertical descent velocity that it will be able to free-fall from, care has to be taken to ensure that the rocket exhaust does not contaminate the surface regolith that is to be collected, and that the rocket does not cause unacceptable levels of erosion to the surface, which could jeopardise the mission. In addition to the work being done in the scope of PhSR, the European Space Agency is funding an experimental facility for investigating these nozzle expansion problems; the current progress of this is described. To support this work, an uncoupled hybrid computa...

Numerical and experimental capabilities for studying rocket plume-regolith interactions

The Orbital 500R is a commercial semi-reusable, two-stage launch system under development by Orbital Access. The focus of this paper is a numerical aerodynamic analysis of the reusable first stage spaceplane, capable of powered and glided flight. The vehicle is intended to release expendable upper stage(s) to inject 500-kilogram of payload into low Earth orbit. Ejection of the upper stage(s) is expected to be above 85-kilometres, after which the first stage spaceplane will perform re-entry and landing. Toolset validation and the establishment of best practice is an integral aspect of assessing and optimising system performance and controllability. Steps have been taken to progress this, related to the use of Fluid Gravity Engineering Navier-Stokes solver, ANITA. The validation activities were focused around the characterisation of the aerodynamics of two wing-body experimental models, tested across a range of angles of attack at Mach 4.0 and 8.2 respectively. The results have provided a sufficient level of confidence in the use of ANITA to assess the aerodynamic performance of spaceplane wing-body configurations for supersonic and hypersonic flow regimes. A computational assessment of the vehicle's aerodynamics has been performed by means of both engineering-based tools and Navier-Stokes CFD computations. Lessons learned from the validation activities fed into the use of ANITA to characterise the vehicle's aerodynamics from Mach 3.0 and above. The lift and drag characteristics of the vehicle were found to be very comparable to those documented for the X-34 reusable launch vehicle from Mach 2.0 up to Mach 6.0. Furthermore, strong code-to-code agreement was observed with the SU2 and ANSYS-Fluent Navier-Stokes solvers for aerodynamic characteristics at Mach 3.0.

Aerodynamic database development for a future reusable space launch vehicle, the Orbital 500R

This paper presents a novel method for multi-objective optimisation under uncertainty developed to study a range of mission trade-offs, and the impact of uncertainties on the evaluation of launch system mission designs. A memetic multi-objective optimisation algorithm, MODHOC, which combines the Direct Finite Elements transcription method with Multi Agent Collaborative Search, is extended to account for model uncertainties. An Unscented Transformation is used to capture the first two statistical moments of the quantities of interest. A quantification model of the uncertainty was developed for the atmospheric model parameters. An optimisation under uncertainty was run for the design of descent trajectories for the Orbital-500R, a commercial semi-reusable, two-stage launch system under development by Orbital Access Ltd.

Robust multi-objective optimisation of a descent guidance strategy for a TSTO spaceplane

For both sample return and insitu-analysis missions it is
vitally important to understand the physical phenomena
related to the propulsion subsystem and terrain interactions
during landing of the spacecraft. The plumes of the retrorockets
will cause erosion and chemical contamination to
the celestial body’s surface which will affect the landing
manoeuvres and measurements of the regolith in the vicinity
of the landing site. Furthermore, the generated regolith dust
clouds and the reflected plume flow can affect the spacecraft
force and torque balance as well as posing problems for the
proximity navigation equipment and sensitive external
surfaces. To investigate the interaction between rocket
engine plume and regolith at a fundamental level, in support
to all planetary and lunar landing environments, such as the
Phobos Sample Return mission, a dedicated facility is
designed, manufactured, and housed at the University of
Glasgow.
The results presented will cover the design calculations
of the vacuum facility with the added mass flow from
pulsating hypersonic nozzles. The methodology behind the
specially designed heat exchanger operating at different
temperatures, pressures, and frequencies representative of
the various landing scenarios is also presented. Analysis to
determine the most suitable material to act as regolith
simulant, due to the variations in Earth’s specific gravity
compared to Martian and airless bodies, is also discussed.

Jim Merrifield

Papers

Vehicle and mission design of a future small payload launcher

Numerical and experimental capabilities for studying rocket plume-regolith interactions

Aerodynamic database development for a future reusable space launch vehicle, the Orbital 500R

Robust multi-objective optimisation of a descent guidance strategy for a TSTO spaceplane

Test Facility to Investigate Plume-Regolith Interactions