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High Power Electric propulsion: a Roadmap for the future

Final Report Summary - HIPER (High power electric propulsion: A roadmap for the future)

Executive summary:

Space is a very harsh environment: temperatures are extreme on both sides, distances are measured in terms of millions of kilometres and vacuum make things even worse. To successfully explore space, there are some basic 'building blocks' that are needed as a minimum:

1. an effective, reliable and cost-effective way to access Space from Earth (which means a launch system)
2. a robust spacecraft
3. a shielding system from cosmic rays and solar flares
4. a power generation system
5. an efficient and reliable propulsion system.

HIPER project was aimed to give a boost to points four and five above. A choice was made to focus exclusively on electric propulsion (EP) systems and on related high power generation, either because of the know-how and expertise of the participating partners and for the intrinsic advantages of EP technology with respect to the traditional, well known, chemical propulsion (namely rocket engines).

For the latter aspect, it is useful to address state of the art specific impulse capabilities, which are at least an order of magnitude higher for EP systems than for chemical propulsion and the related overall efficiency. On the other side, chemical do perform better -at least for now- in terms of thrust and of reliability and do require less electric power than EP.

HIPER aimed to advance knowledge and technology one step further with respect to the original state of the art, either for the selected EP systems or for the related power generation systems.

As a starting point, a study has been conducted on a set of missions that could be operated with high power EP systems. Although space science and exploration will be the main driving force behind the technological development of EP and high power generation systems, novel techniques and methodologies will equally benefit commercial and utilitarian space missions, thus generating a very significant impact on Europe's capabilities to access and exploit space.

HIPER not only addressed technologies, but also attempted to consider major technological efforts in the framework of social and political scenarios, both internal to Europe and with respect to non-European partners. In this respect, there have been discussions and meetings with National Aeronautics and space Administration (NASA) and large United States of America (USA) companies to share HIPER results and harmonise future roadmaps.

A number of results have been obtained in the project's timeframe, including the definition of a set of exploration and transportation scenarios; the design, prototyping and initial test of a high power (20 kW) hall effect thruster (HET), the design of a multi-gridded Ion Engine and the development of a high current cathode, the design, prototyping and initial test of two 100 kW class magneto-plasma-dynamic thrusters (MPDT); the design, prototyping and initial test of an inflatable SA with Fresnel lens concentration system and the feasibility study for a 200 kW nuclear reactor power generation system in space.

Project context and objectives:

HIPER project was conceived to depict a roadmap for future space exploration and transportation, which could benefit from technological advancements in EP technologies and from the related advancements in power generation in space.

Rationale for such ambitious goals rely on the intrinsic efficiency of EP on one side and on its proven heritage since the late 1960s in a number of scientific and commercial missions and on the fact that high power EP needs high power generators onboard, on the other side.

To achieve the goal an interdisciplinary team of 20 partners from six European Union (EU) countries, has been formed, grouping most of the specialists in EP and in power generation in space.

The whole HIPER project has been divided into five main technical work packages (WPs), each of them fully dedicated to a different topic. The first two technical WPs are respectively devoted to mission analysis and to electric power generation. The other three WPs are instead focussed on three different EP concepts.

Mission analysis, propulsion system requirements and recommendations WP

The main goal of this WP was the definition of future scenarios for European space transportation and exploration and of ensuing possible missions' objectives. The reference scenarios have been built by taking into account technical as well as political challenges and the parallel evolution of non-European trends. Definition of requirements for power generation and EP systems, based on the above mentioned scenarios, have been produced as input for the other WPs. Medium term scenarios as well as more futuristic, long term scenarios have been produced and presented to various EU and international players.

In-space power generation WP

The goal was to design a power generation subsystem capable of providing power levels ranging from tens of kW to thousands of kW. SAs technologies and nuclear reactors were studied. Both solutions have been considered and thoroughly investigated. A prototype of an advanced SA based on concentration has been manufactured and assembled, whereas the nuclear reactor option has been preliminarily assessed and seized in terms of mass, volumes, thermal control systems and shielding from radiations.

High power and hall thruster development WP

Within this WP the design, manufacturing and development of a HET prototype working at a power level of 20 kW has been faced. A very preliminary experimental campaign aimed at measuring the thruster performance and to validate technological solutions envisaged during the thruster development has been conducted at the end of the WP activities. It has to be noted that the produced prototype is the most powerful HET produced in Europe.

High power gridded ion engine development WP

This part of the project was aimed at investigating novel concepts for thruster grid choice in order to design gridded ion engines (GIEs) efficiently operating at higher power levels. The very novel approach of dual stage, multiple grids engine has been analytically investigated and grids number and distance optimised for future development. At the same time, design, development and experimental campaign on high current hollow cathodes to be used as neutralizers for this class of thrusters has been conducted.

MPDT development WP

Within this WP the design, manufacturing and test of two prototypes of MPDTs, one operating with an externally applied magnetic field and one operating with the sole self-induce magnetic field, has been conducted. Both thrusters have been successfully operated during preliminary tests in vacuum chambers; in parallel the first European multi-channel hollow cathode has been developed and successfully tested for more than 100 hours.

Considering the potential impact of the EP in a near-mid future term, the work carried out along the three years of HIPER program has a significant strategic importance. For the first time a large group of private companies, enterprises, research centres and universities has been put together to actively discuss about the necessary steps to be taken in order to define a roadmap for the future development and exploitation of high power EP. And the effort could have hardly been more successful, resulting in a fruitful cooperation among all the involved partners, who worked together under a common objective.

Project results:

Introduction and major benefits of EP

In HIPER programme, the analysis of possible mission scenarios is an essential prerequisite to carry out an effective road mapping and long term technology planning study. Building a credible and motivated long term reference scenarios is however a task which is very complex and implies considerations much beyond the purely technical aspects. Political, economic and social constraints all influenced the evolution of space policies over many decades and many of these elements may undergo unpredictable changes.

EP can play a very important role in future space exploration programmes by enabling more affordable and sustainable space to space missions. Modern EP technologies enable mass saving, launch flexibility, long interplanetary journeys and faster missions with no gravity assist constraints. This also opens the way to transferring large payloads through the solar system more affordable than in the past. Besides, larger payloads can be transported by increasing the operational power level of the propulsion systems.

From the economical point of view, high power EP applications will have a significant reduction on the overall cost of access to space by improving the near Earth orbit transfer and maintenance cost effectiveness of institutional and commercial spacecraft.

With respect to long term sustainability, high power EP will provide the only sustainable means of transportation in the near-Earth region and across the solar system capable of allowing effective exploitation of space resources and environment.

Target mission scenarios

Selected scenarios try to harmonize the long term plans by the main space agencies, with the 'sensitivity' and prediction capability of experienced space executives and with the visionary inputs by small companies which are at the leading edge of space propulsion innovation. HIPER mission analysis team was devoted to define some near and long term mission and transportation scenarios which could specially benefit from high power EP. The HIPER list of target classes of missions includes:

1. orbit transfer in the Earth-Moon system
2. robotic Mars sample return (MSR)
3. near Earth objects (NEOs) exploration, exploitation and risk mitigation
4. Mars and its moons (Deimos and Phobos) fly-bys and science
5. outer solar system and beyond robotic exploration and science.

Mission analysis

Different strategies have been adopted to analyse the mission profiles previously described, since, even if we are always considering low-thrust transfers, different targets requires different manoeuvres.

For what concern missions taking place within the Moon-Earth system, a circular restricted three body model (CR3BP) approach has been used to study the problem. This is the basic model to study the motion under more than a single gravity field and is based on the assumption that the moving body (m3) has a mass so small that it cannot influence the motion of the two massive bodies (the primaries, m1 and m2) exerting the gravitational influence. It can be shown that, under these assumptions, there are five positions of equilibrium in space, the Lagrangian libration points Li, where the gravitational fields of the primaries combined with their centrifugal force are in balance.

For Mars Sample Return, the whole analysis has been divided in several parts. Initially there is an Earth-escape phase, where the spacecraft spirals out of earth sphere of influence; then there is the interplanetary travel between Earth and Mars sphere of influence, followed by a Mars-capture phase. After a certain time of stay, the spacecraft starts its travel back, again divided in three phases: Mars-escape, Mars-to-Earth interplanetary transfer, Earth re-entry. Two different analyses have been carried out for interplanetary phases, with the goals of minimising the fuel consumption and transfer time.

The mining scenario instead calls for a mission which travels forth and back between Earth and the target asteroid. Starting in a Lagrangian point of the Earth-Moon system (either L1 or L2) the spacecraft travels interplanetary to the asteroid where it is orbit the small body. After some stay time to load the mining material, either raw or processed, the spacecraft returns to the Earth. This mission consists of two interplanetary trajectories. Other manoeuvres like capturing at asteroid are neglected for the activities presented in this document. In fact, the target asteroid can be very small and the delta-v for the orbit insertion is likely to be negligible. Besides, another assumption for this kind of mission is that the spacecraft is powered by a nuclear reactor. Thus, constant power is available for the EP system independent of the distance from the Sun.

EP subsystem main requirements and key recommendations

In parallel to the mission scenarios definition phase, a preliminary definition of requirements for power generation and EP has been carried out, in order to provide a solid basis for the assumptions used in the preliminary mission analysis. Preliminary requirements are set for each voltage level used to accelerate ions.

As the high power solar EP (SEP) is considered a technology available in the near/medium-term, only the high power HET and GIE technologies have been considered as possible developments in this timeframe. Therefore in SEP case, only requirements for medium and high voltage levels are drafted. On the other hand, in nuclear EP (NEP) case also the high power MPDT technology has been taken into account as possible development in the long-term.

In-space high power generation

Solar Power Generation is one of the two available options (together with nuclear power) to generate a high amount of electric power in space. Several topics have been investigated with the purpose of making a step forward in solar power generation technology: light concentration system design, inflatable substrate and flexible electrical network structural design, prototyping and verification. The solar cell trades started from the very latest achievements in crystalline solar cells. Based on that, a component suitable for light concentration has been identified. This component has been integrated on top of a flexible structure in combination with the concentrating lenses and with an innovative 'smart' deployment mechanic based, also, on shape memory alloys (SMA) performances. Power management and distribution (PMAD) is a critical aspect of the whole system design which can undermine all the other technical advances. A direct drive approach has been selected as the most suitable solution and an architecture based on solar direct current (DC) voltage level of about 300 V and a current level of about 800 A has been extensively simulated.


As stated above, the first requisite is that the power system has to deliver a total power of 250 kW and this power is generated by a solar array (SA). In the present study an 'intermediate' approach, called almost direct drive architecture (ADDA) is presented. Such architecture is, possibly, the best trade off in terms of performance versus weight and compactness of realisation.

The proposed system is based on the following assumptions:

1. total power required by the thrusters is 250 kW
2. the SA is composed of P sections (with e.g. P=10)
3. the open voltage of each wing in the SA generator is approximately 200 V
4. there are P power subsystems working in parallel, each transferring 250/P kW to the main bus
5. when considering architectures with regulated main bus, the bus is at 300 V
6. the subsystems are 'star' connected to the 'cold' node of the main bus
7. there are T thrusters (with, e.g. T=10), each rating 250/T kW all connected to the 'cold' node

In some power system architectures considered here there can be a battery connected to the main bus through a battery discharge regulator (BDR). The battery delivers energy during possible fast power fluctuations introduced by the thrusters. Battery discharge has very strict time limits due to the high power levels involved. The battery introduces a delay in the shut-down of the power processing unit (PPU) or, better said, it allows to quickly reducing the power absorbed by the thrusters without introducing PPU instability or even causing an unwanted total shut-down of the PPU. The use of a maximum power point tracker is essential for the mission due to the need of maximising the efficiency in different illumination conditions. Furthermore, given the extremely large dimensions of the SA, a fractioning of the total SA power over several wings is a desirable feature.

As already anticipated, the result of the trade-off analysis is ADDA. As for the other architectures the total power absorbed by the thruster is partitioned among wings and sections of SA. We considered a SA and its P sections. At first we assume that a single SA section is connected to a thruster of equivalent power. All P sections but one are connected in parallel, the P-1 sections are directly connected to the load. We firstly assume that the thruster works is a 'constant resistance' condition, that is its equivalent resistance and it is kept constant independently from the electrical operating point of the of the P-1 sections of SA. This means that power absorbed by the thruster can be increased by simply increasing the voltage of the direct drive bus. In the sequel the load resistance is varied, so that motor on/off situations and power fluctuations/variations are simulated.

Solar cells with light concentrators

As anticipated at the beginning of this section, the likeliest evolution of the space borne solar cell assemblies is towards thinner and lighter structures, capable to be used in combination with solar generator systems having a reduced volume at launch, combined with outstanding electrical performances.

The solar cells which have been be used in the frame of bread boarding activity of HIPER are multi-junction; their peak in efficiency is designed to be reached at concentrating factors of decades, but they will be also suitable to work at levels around six to eight times the AM0 intensity.

The light concentrator is a Fresnel lens. A lens with an axial symmetry will focus sunlight onto a straight line, centred over the active area underneath. A deployable sustain will position the lens over the solar cell line. Each lens operates coupled to a string of 10 solar cells and it is realised by using space grade silicone or an optical grade plastic (e.g. polycarbonate). The Fresnel lens is comprised of precisely formed ridges which refract the light from a 10 cm aperture down to a strip of light focussed in the middle of a 1 cm wide solar cell, to leave margin for pointing error. Lens positioning will follow the deployment of the panel/substrate, so that the subsystem lens-sustaining structure will initially be packed. The SMA adopted in the frame of HIPER solar panel will be an alloy of nickel and titanium.

SA architecture and prototyping

The design of the SA module is as follows:

1. an inflatable structure has been thought to tension the membranes which sustain the solar cells and his architecture allows to deploy a huge surface
2. the substrate is composed by a set of modules all equal to each other
3. each module is made by a layer of kapton
4. each module is rigidised by a sustaining frame
5. the modules are joint by passive hinges
6. the deployment relies both on passive hinges and SMA
7. printed circuit lines are used to connect electrical items.

To demonstrate the functionality of the deployment structure, a breadboard has been built. It is made of three modules each equipped with dummy lens and a set of electrically representative items.

Nuclear power generation

HIPER team investigated nuclear fission electrical power generation for the exploration of the outer solar system. Mission analysis identified a range of applications from one way journeys to Uranus, return missions to the Jovian and Saturn planetary systems and multiple, shorter infrastructure, or manned, delivery. The nuclear 'space tug' concept was seen as the best fit for these applications.

The objective was to 'develop a roadmap for the longer term provision of nuclear power sources for space EP supported by critical risk simulation and modelling'. The technical starting point was a Rolls Royce nuclear technologies for space applications survey and an Acta shield design study which provided the baselines for modelling and simulation. The three main pillars of the investigation were the core reactor, the shield and the power conversion system. However the potential to achieve the individual capabilities had also to be consistent with a viable, overall system architecture as well as each other.

Analysis of previous studies showed specific mass to be the principal design driver and a specific mass of 25 kg/kWe was thought the best that could be achieved with current and emerging technology. In principle this constrained the power generating capability to 200 kWe which became the target for a concept design.

Initial investigation showed Brayton cycle power conversion to have the most competitive specific mass for the 200 kWe power range. No clear advantage was found between the relative merits of direct cycle, gas cooled, epithermal and indirect cycle liquid metal cooled fast reactor based systems. Consequently the Roadmap identifies the technical development required to realise the concept designs for both technologies developed during the study.

Study architecture

The study was in three stages. A comprehensive survey was made of previous projects and studies to establish realistic design targets. Achievable performance was then investigated through the modelling and simulation of reactor core, shield and power conversion design options. The results provided the basis for a fission nuclear power generator concept design together with a roadmap for the technical developments required for its implementation.

Reactor core physics

Both Indirect liquid metal cooled and direct gas cooled reactors could generate 200 kWe over the 10 year lifetime or longer if required. Fuel composition and reactivity control are compatible with water immersion safety requirements.

The more compact liquid metal Indirect Brayton cycle reactor was 25 % of the mass of the direct Brayton cycle (476 compared to 2 017 kg) before consideration of the mass of a heat exchanger to transfer the thermal energy from the liquid metal to the Brayton cycle operating gas. The direct cycle core physics is dominated by the inlet/exit gas flow paths, which occupy about 35 % of the fuelled region volume in the design which was analysed.

Risk analysis indicated scope for technical development to target a higher gas pressure, accepting increased core resistance. The gas flow areas could then be halved, with a 30 % reduction in the number of fuel bed annuli, to give a 10 % reduction in core size. Neutron and gamma escape fluxes from the core were calculated at all exterior surfaces. Both showed radial distributions. Activation of the coolant gas was also calculated for the direct design and does not pose problems.

Both reactor designs considered in the risk analysis exhibited quite large excess reactivities with the control media 'withdrawn'. These were left in the concept design to provide margin for loss of reactivity arising from the implementation of engineering features and realistic materials. These margins could be 'trimmed' in future iterations of the designs, to reduce the fuel loading and/or fuel enrichment.

Radiation shielding

The shield modelling and simulation was based on a simple spacecraft configuration using US Government Monte Carlo N-particle extended (MCNP-MCNPX) code. For the direct cycle the neutron and gamma flux radiated from the reactor is seen to decrease as one moves from the centreline thus allowing a corresponding reduction in shield thickness toward the outer edges. The Indirect cycle had a similar distribution with higher radiation intensity. Shielding requirements were based on the US SP100 project criteria. The materials are selected and arranged to optimise performance for radiation shielding, structural and thermal resilience. Coolant pipes can be routed round the shield and control rod penetration leakage would be minimised in detailed design.

The final modelling results gave an end of life gamma dose or neutron flux generally lower by a factor of two to four than the requirement apart from the gamma dose at payload entrance for the Direct cycle reactor, where the margin is only of about 20 %. Control rod design minimising the distance between reactor core and shield gave a 50 % mass saving.

Power conversion

The power conversion options investigation showed the Brayton cycle to be more efficient than a thermionic or thermoelectric design, with cycle efficiencies of 17 % to 19 % compared with 5 % for the latter. Specifically the indirect cycle was the more efficient as a result of a reheat loop but it is a mechanically more complex.

The closed loop Brayton cycle has radial turbines and compressors and a helium and xenon operating gas. Efficiency is increased by a recuperator and by minimising the pressure drop over the system, especially in the radiator. In the indirect cycle a lithium liquid metal coolant is heated by the core which in turn heats the operating gas in a heat exchanger. For the direct cycle the operating gas is heated as it flows through the reactor core. The lowest radiator mass was achieved with a two-staged compressor and free power turbine arrangement.

Modelling and simulation showed the indirect cycle to be more efficient than the direct cycle, with an 11 % lower radiator mass. The optimum turbo-machinery configuration was a two-shaft compressor with a reheat loop driving a free power turbine. The reheat loop increases the temperature of the fluid entering the free power turbine and consequently the temperature of the fluid at the radiator inlet. The heat exchanger for the indirect cycle was not thoroughly investigated as it was not seen as being the main limiting component of the power conversion unit. However, it is anticipated that any heat exchanger design will also be limited by the creep behaviour and the materials employed to achieve a ten year mission life and this will constrain the maximum temperature of the operating gas.

The modelling and simulation identified two critical mass design drivers, namely radiator size and turbine inlet (determined by core outlet) temperature. Technical developments to raise turbine inlet temperature are:

1. turbine inlet design. Realise the benefit of the turbine rotor experiencing a lower stagnation temperature (around 80 to 200 K) because of its rotation.
2. turbine blade cooling. Bleed cooler coolant gas from the compressor onto the turbine rotor can keep turbine blades up to 200 K below the inlet gas temperature.
3. refractory metal alloys. High temperature refractory metal alloys such as niobium which is not susceptible to oxidation operating in with a closed xenon/helium cycle.
4. ceramic materials. Ceramic materials have the thermal and creep properties for operation up to 1500 K but need to develop resilience to stress fracture.

Singly or in combination these enhancements open the way to core outlet temperatures in the range 1200/100 K (moderate risk) to 1500 K (high risk).

High power HET

HET, also called stationary plasma thruster or closed electron drift thruster, is an advanced propulsion device that uses an electric discharge to ionize and accelerate a propellant gas. The basic concept of a HET was suggested in the early 1960s almost simultaneously in the former Union of Soviet Socialist Republics (USSR) and in the USA. However, it is only in 1972, that the first demonstration was given in flight by the soviet satellite Meteor. Since the time of pioneer works, several hundreds of spacecrafts have flown with hall thrusters.

The recent success of the SMART-1 lunar orbiter mission of the European Space Agency (ESA) demonstrates the possibility to employ HET as the main propulsion means for an interplanetary journey. Currently, 1 to 2 kW-class HET are employed for station keeping and attitude control of geostationary communication satellites. Nevertheless, ambitious robotic missions like exploration of outer planets of the solar system and far-off comets as well as transfer of cargo vehicle to support crewed missions require very high power EP systems.

The development of high-power HET and of any large-scale device represents certain design and technological challenges. The goal of the HIPER WP4 was to progress in the field of high power HET for exploration missions. Six partners from the industry and scientific organisations gathered to address the different technological challenges related to high power HETs: Alta (Italy), IPPLM (Poland), CNRS (France), Tecnalia (Spain), Onera (France) and Snecma, Safran Group (France). Our approach was split into three main steps:

1. Assessments of the main technological options for the development of high power HETs. The manufacturing of a high power HET prototype needs to consider with specific justifications the material and thrusters pieces interacting with the high energetic plasma flow in the thrusters. A first innovation to be met is the manufacturing of a ceramic discharge chamber of large diameter (about 300 mm). The major limitation was due to the manufacturing process. An alternative approach was developed by the brazing ceramic sectors of the discharge chamber. Another addressed issue was the interaction of the plasma with the discharge chamber. It is well known that the ion and electron flows of the plasma will interact strongly with the ceramic discharge channel and will affect the thruster performances. The material characteristics of the ceramic are then to be quantified and especially the erosion rate due to energetic ions and the emission yield of secondary electrons, which are critical ceramic properties. The obtained results highlight that the SEE measured on channel material that has not endured the specific HET environment could be very different from that of the same material under HET working. To be somewhat more representative, the secondary electron emission (SEE) must be measured on materials that have been aged under both ion and electron irradiation.
2. Preliminary design of a high power HET laboratory model. Following the previous studies, it was proposed to manufacture a 20 kW Hall thruster. Several topics were considered for the design of a high power HET: the thruster architecture, its thermal optimisation, the design of the electromagnet, the implementation of a cathode delivering up to 70 A and the modelling of the plasma in the discharge channel. A monolithic HET architecture was proposed. In order to offer more thermal margin, a splitting of the neutral injection and anode was proposed. The magnetic field of the HET was generated by coils and a ferromagnetic circuit. The magnetic field of the thruster was also measured and a very good agreement with the computations was concluded. Concerning the cathode, alternative technologies were studied. Nevertheless, due to costs and time limitation, it was decided to implement a state of the art hollow cathode. Finally, one-dimensional and two-dimensional modelling of the plasma in the discharge channel of the 20 kW thruster was carried out, the design of the laboratory model was set and the drawings to the thruster released.
3. Manufacturing, assembly and testing of a 20 kW HET laboratory model. Following the two previous steps, the last year of this activities was devoted to the manufacturing of a 20 kW laboratory model and its testing in a vacuum facility in order to validate the technological choices made in the HIPER project. The thruster was tested in a cryogenically pumped cylindrical vacuum chamber called Pivoine. As this test facility was designed for the testing of low power HETs, an upgrading was realised before firing the PPS-20k ML. The xenon feed system and the electrical hardware were modified in order to perform high power tests. Another important modification of the Pivoine test facility was the design and the manufacturing of a new test balance. After ignition of the thruster, preliminary test were realised in order to check the behaviour of the new thrust balance and cathode and to validate the test facility upgrading. The thruster was then tested at high power. A total number of 30 operating points were characterised. The 20 kW thruster could be operated very successfully and reached the expected performance, in accordance with the specified specifications.

As was previously mentioned, the PPS-20k ML HET is a dismountable thruster. Neither complete parametric studies nor magnetic optimisations could be performed during this first test campaign. Consequently, possible improvements of the thruster performance deserve future test campaigns.

It is also important to notice that propellant selection is of critical importance. Xenon was selected as the propellant during our test campaign because it is the only propellant under utilisation in the current and near term missions employing Hall thruster propulsion. Potential alternative fuels for Hall thruster must be addressed in order to reach economical implementation of high power propulsion.

In conclusion, thanks to the HIPER project, the successful implementation of high power Hall thruster was demonstrated following the design, manufacturing and testing of a 20 kW laboratory model. Evaluation of high power Hall thrusters with alternative propellants is suggested in the future in order to offer an economical access to in-space propulsion.

High power GIE

During the course of the WP5 of the HIPER study a novel type of GIE called the dual stage GIE (DS3G) has been studied. The study focussed on both the DS3G thruster concept and on the discharge hollow cathode needed for the most demanding HIPER mission requirements.

DS3G thruster

A DS3G is a gridded ion engine where instead of using a single ion optics stage to extract and accelerate an ion beam, two stages are used to separate the extraction and acceleration process. The main effect of this separation is the fact that, unlike conventional GIEs, the thrust density of this novel thruster (up to 12 N/m2 at an Isp f 10 000s have been simulated) can be increased while also increasing the Isp up to values that are several time higher than those currently produced by conventional GIEs. This results in a compact thruster that can process tens of kW of power producing high efficiencies.

To realise a separate acceleration and extraction stage an extra grid must be used hence resulting in a three-gridded GIE. The first aspect of a DS3G that has been analysed is it applicability and the operating conditions under which its thrust density is in excess of those produced by conventional GIE.

Using analytical model available in the literature and numerical simulations performed with the FFX code, the trend of the thrust density provided by a DS3G in comparison to that provided by a conventional GIE has been derived as a function of the ratio between the voltage drop applied to the acceleration stage to that applied to the extraction stage.

As it can be seen the DS3G offers performances in excess of those of the GIE only if a value of higher than 0.6 is used. Assuming that the DS3G is obtained adding an extra grid to a GIE able to deliver an Isp of 4 500s this means that the DS3G will provide an increase in thrust density only for specific impulse level in excess of 6 200s. Numerical simulation showed that at 10 000s the DS3G was able to successfully extract the same beamlet current as a GIE working at 4 500s delivering slightly more than twice the thrust density.

To deliver high specific impulse levels, high voltages must be employed on the thruster screen grid. The presence of such high voltages produces concern regarding the thruster lifetime since the charge exchange ions (CEX) created within the thrust ion optics can impact on the grid with high energies hence producing high erosion rate and consequently short grid lifetime.

The DS3G was found to be able to deliver lower beam divergences (about 40 % less) than the RefGIE and a much higher thrust density. The DS3G also suffered from lower erosion rates than the RefGIE and consequently was able to provide a longer lifetime. The DS3G lifetime was found to be about 45 000 hours whereas the one of the RefGIE 37 000 h.

DS3G discharge hollow cathode

Besides analytical and modelling work on the DS3G thruster, a hollow cathode able to deliver the discharge current needed by a DS3G has also been designed and tested. The cathode design has been carried out for the highest current case between those obtained elaborating the mission requirements from WP2.

A cathode able to deliver a current of 180 A has been designed and tested. The design has been carried out taking into account that for the NEP mission to Jupiter a lifetime of 17 000 hours must be delivered. The cathode has then been built by QinetiQ according to the specifications above whereas the rest of the experimental setup has been built at the University of Southampton.

The cathode insert has been equipped with five thermocouples to monitor the temperature profile. The cathode has also initially been manufacture with a smaller 3 mm diameter orifice. The main goal of the experimental campaign was to gather more information about the functioning of the cathode and the influence of the orifice size over the cathode insert temperature distribution.

The cathode was tested with three different orifice sizes (3, 5 and 8 mm), at various current levels (ranging from 25 to 180 A) and with various xenon mass flow rates for a total of 130 h of which 30 were at the 180A discharge current level.

The maximum current of 180 A was achieved with an 8mm orifice diameter using 10 sccm of xenon. The insert temperature distribution at 180 A had an average value of 1 380 °C and showed a strong peak at the insert downstream end of 1 460 °C.

After the test, the cathode was visually inspected and some of the parts of the experimental setup were analysed using energy dispersive x-ray (EDX) spectroscopy. The cathode did not show major signs of wear and the EDX spectroscopy found deposition of copper (coming from the anode support structure) on the external surface of the keeper and of tantalum and barium (coming respectively from the orifice plate and from the insert low work function compounds) on the inner surface.

The lifetime of the cathode has then been extrapolated using the measured insert temperature profiles, by applying three different life time criteria.

According to the first criterion the cathode will able to deliver up to 20 000 h at 180 A assuming the worst case for which all the insert is at the peak temperature of 1 480 °C whereas according to the other two criterion the lifetime at 180 A will be 15 000 and 3 000 hours, respectively. At the 150 A, the cathode lifetime will be in excess of the required 17 000 h according to all three criteria.

Considering the different answers coming from the lifetime evaluation criteria a modification of the cathode design has been proposed to assure that the cathode will be able of deliver the required lifetime.

The numerical investigation of the dual stage ion engine showed that it is able to provide higher thrust densities, lower jet divergences and longer lifetimes (in excess of 100 000 hours if graphite grids are employed) than conventional GIEs.

This makes the DS3G the best candidate for the high specific impulse missions studied in HIPER and for all future high specific impulse missions. Future work will focus on building a DS3G prototype to verify the numerical predictions.

For what concerns the hollow cathode experimental campaign the test will be considered successful mainly considering the limitation in time and budget and the recent efforts performed by other institutions to build a long lifetime cathode able of providing 250 A using LaB6 insert in conjunction with water cooled anode.

Moreover the cathode was tested in diode mode and this test setup has been found to lead to higher insert temperatures than those relative to cathode operation inside a discharge chamber.

The development of such high current cathode is not only relevant to DS3G but also important for the future development of high power HETs studied in HIPER given their need for a long lifetime and high current neutralizer cathode.

Further work should include long term testing (with temperature measurements) with a cathode modified according to the design recommendations herein and the development of a full numerical model of the cathode plasma flow and a thermal model.

High power MPDT

MPDTs have long held the promise of high exhaust velocity (c) at MW power levels. The combination of high c and high power in a compact device is especially beneficial for demanding missions such as the human exploration of other planets, which will require lightweight, high power density propulsion to be feasible. Nevertheless, despite more than 50 years of studies, MPDTs have somehow failed to meet researchers' expectations entailing a slow but constant decline of interest in such a technology. At present, MPDT research is still at a fundamental rather than a development level. Research goals include increasing in thrust efficiency and operative lifetime. In particular applied-field MPDTs appear to be more amenable to near-term applications since they retain a satisfactory thrust efficiency (25 to 30 %) even at power levels in the range of 100 to 200 kW. Nevertheless, the ability to fully realise the transportation and economic benefits deriving from application of high power EP is strongly dependent upon the development of suitable electrical power sources. At present, these are essentially based on photovoltaic SAs, although some early attempts at small radio-isotope and fission-reactor EP missions to the outer planets are under study. These new trends will allow MPDTs to be re-considered as viable propulsive option in the near future.

In this context, the efforts undertaken by Alta and IRS during were aimed at recovering and broadening the past knowledge concerning MPDT technology as well as at paving the way for an advanced European MPDT to be developed and test in future programs.

Taking advantage of their different backgrounds, Alta focussed on a pulsed, quasi-steady device whereas IRS developed a steady-state MPDT. Both of them operated at about 100 kW using argon propellants. At the same time, Alta developed and tested high power cathodes to assess the effectiveness of the multichannel technology as a viable option for future EP devices.

Activities at Alta

Alta MPDT was designed taking into account the most promising design options from the past experiences and findings. A coaxial configuration with a central multichannel hollow cathode and a flared anode was finally selected.

Measurements were obtained for a variety of currents, mass flow rates and magnetic field strengths in a power range between 20 and 250 kW. Tests were carried out in Alta's IV-10 vacuum facility. With a volume of about 200 m3, IV-10 allowed for current-pulse duration up to 1 s maintaining a back pressure lower than 3.10 to 4 mbar as well as for the minimisation of the environmental interaction with the plume. Although the shot duration was still too short to achieve steady-state thermal conditions, it allows for direct, time resolved thrust measurements. To this purpose a new single-axis thrust stand was designed to improve the full scale and the frequency response of the existing thrust stands commonly employed for high power devices. A maximum thrust efficiency of about 30 % was obtained at 200 kW for an applied field of 120 mT and a mass flow rate of 60 mg/s. At 100 kW, for the same mass flow rate and magnetic field, the thruster reached thrust efficiency slightly higher than 20 % and a specific impulse of about 2 500 s.

These results appear very promising since the thruster performance parameters (specific impulse, thrust efficiency and power-to-thrust ratio) are among the highest ever measured for argon-fed MPDT in the power range investigated. Besides, the experimental set up has proven to be reliable and adequate to perform long-lasting (500 ms), quasi-steady pulses allowing for time-resolved, highly accurate thrust measurements.

To assess the cathode erosion process during steady-state operations, a reduced-scale multichannel hollow cathode was tested for 100 hours. The cathode wall temperature along the axis was measured by using an optical pyrometer since temperatures well in excess of 2 000 °C are needed to sustain a stable, diffuse, discharge. In this condition, the material evaporation is the main erosion mechanism. It was found that the start-up phase is characterised by localised arc attachment on the cathode surface where the temperatures at the arc root are often higher than the tungsten melting point (about 3 400 °C) leading to erosion rates as much as three to four orders of magnitude higher than that seen in steady-state operation (around 1 ng/C) even for extended operation (the ratio between the steady state time over the transient time was about 120).

The McHc assembly (rods pattern and their location with respect to the exit section) was found to highly influences the erosion rate. Unfortunately the reproducibility of the McHc manufacturing is questionable and remarkable steps in this direction are still to be taken.

Activities at IRS, Stuttgart

At IRS a steady-state, AF-MPDT, called ZT-1, was design and successfully tested at 6 kW. The data gathered during the experimental test campaign on the ZT-1 allowed for the identification of a guideline towards the design of a 100 kW, steady-state device, named SX-3.

Moreover, the Alta's hollow cathode was also tested at IRS to include spectroscopic analysis. Unfortunately, the test carried out at IRS appeared to be problematic since the hollow cathode was already depleted due to the hours of operations at Alta. Nevertheless, the metallographic analysis undertaken at IRS revealed interesting behaviours of the tungsten under continuous thermal stress. These findings will allow for an improved design procedure in future activities.

European AF-MPDT

If MPDTs are to have a role in future EP missions, completely radiative-cooled versions must be developed with lifetime requirements not lower than 2 000 to 3 000 hours. In this context, the main output of the program is the design of a steady-state, gas-fed, applied field MPDT conceived to include the most promising design options and material selections suggested so far by the main research institutes with the aim of performance and lifetime enhancement.

The thruster consists in a central lanthanum hexaboride cathode and a concentric conical anode. Argon propellant is injected through an orifice obtained in a De-Laval nozzle which is part of the gas feeding system. The overall cathode length from the orifice to the insert was defined by the need of heat rejection. The thruster overall dimensions are 190 mm diameter and 420 mm length and the estimated total mass is about 25 kg without solenoid, cables and bolts. The thruster is completely dismountable allowing for separate electrodes removal and quick inspections. Sealing between the anode and the housing is accomplished by using graphite gaskets whereas the recovery of the clearances due to the thermal expansion is achieved by using a commercial compression spring. The present design suggests the use of a lanthanum hexaboride emitter cathode. The major reason for using LaB6 cathodes is the robustness, high current density and long life exhibited by these emitters during the long years of qualified space operations in SPT Hall thrusters as well as ion thrusters.

As regards the technology road map, two development paths were identified. Since the major technical obstacle for ground testing of MPDTs is the inability of the state-of-the-art vacuum systems to handle the tremendous pumping speeds required for these high power devices, two different paths must be foreseen according to the future availability of testing facilities for steady-state operations.

Path A will allow for a full-scale lifetest opening the possibility of investigating all the relevant technological and scientifically issues at once. By contrary, path B focuses on the investigation of the most critical sub-components, i.e. cathode and magnet. Separate lifetime tests are suggested in order to assess the most probable scenarios at de-rated power conditions.

Potential impact:

HIPER project has shown that the adoption of high power EP systems could represent a turning point for enabling a whole new class of missions, which otherwise would be impractical.

A roadmap is here sketched, considering the current development status of thrusters and power generation technologies and taking into account all the open issues that have to be tackled in the next years, in order to pave the way for a successful and widespread usage of high power EP in space.

For what concerns the thrusters, a substantial amount of work has been carried out in the last three years with the aim of assessing and improving the existing EP technologies that are considered suitable for being used at high power levels (higher than 20 kW).

Among them, Hall thrusters technology is presently the most mature one. Such thrusters have already proven their effectiveness in several space missions, although they have always been designed to operate at lower power levels with respect to ones considered in HIPER framework (say less than 2 kW). However, Hall thrusters are more easily scalable to higher power levels than other propulsion devices and they improve their efficiency in a non-negligible way when the operating power level goes up. A 20 kW prototype has been designed, manufactured and tested in HIPER program, showing an excellent performance. If used in cluster configuration of two, four or eight units, it is possible to reach total power levels comparable to the ones assumed for the previously mentioned space missions (with the notable advantage of having multiple thruster units onboard, which mitigates the risks connected to a malfunction of one of them).

High power GIEs also have a high level of maturity and, like Hall thrusters, at lower power levels GIEs have been widely employed in many space missions. In HIPER project the main task was to focus the attention on a couple of key issues: the cathode development and the definition of an optimal thruster configuration for high power operation (so called DS3G, dual stage three grids). The main strengths of GIEs are the very high specific impulse (a feature that can specially suits the needs of some class of missions) and their extremely long lifetime (which is estimated to be of the order of one hundred thousand hours according to the simulations and the experiments carried out up to now; even if this prediction turned out to be excessively optimistic, the lifetime is expected to be fully satisfying anyway).

The last kind of electric thruster suitable for high power operation is the MPDT. An intensive experimental activity has been carried out on MPDTs under HIPER project, in order to assess their performance and to make a step forward in understanding the intrinsic problems (linked to plasma discharge stability) that have always plagued this class of thrusters, severely limiting their efficiency. MPDTs are expected to be the most effective devices when operating at very high power levels, i.e. greater than 100 kW (up to a few MW). Actually, they are the only EP thrusters capable of processing so much power in a compact size, this resulting in a much higher thrust density (N per m2 of exhaust section area) with respect to Hall thrusters and GIEs. For this reason, going on with a thorough investigation of their behaviour is crucial to further improve their performance and to enable their use for real missions. Of course, considering their current technology readiness Level and the required power level for a single thruster unit, MPDTs will probably play a significant role for space exploration only in the mid-long term.

On the side of the power generation technologies, two are the viable options that have been investigated: solar power and nuclear power.

Obtaining the necessary power from SAs is the most consolidated solution for present spacecraft. Solar energy is readily available in space and the technology to convert it into electric energy is mature and reliable. In the short term, solar power is the preferred choice. Of course there are two main drawbacks. Firstly, if plenty of power is necessary (say hundreds of kW) the size of the solar panels grows larger and larger and handling them becomes truly challenging. Secondly, if the target mission is heading towards the outer planets, solar energy decreases fast with the increasing distance from the sun. Using solar energy beyond Mars, especially for feeding high power devices, is not advisable. Under HIPER program, newly designed solar cells equipped with light concentrators have been developed and assembled in an array as a prototype technology demonstrator. It has been shown that, relying on this technology, it is possible to unfold a group of flexible SAs capable of providing a power of 250 kW around the Earth with a power density of 300 W per square meter and with a specific power of about 150 W per Kg.

As stated above, if even higher power levels are required for a mission or if the spacecraft is operating far from the Sun (further than the Earth), solar energy is not a satisfying solution. In such cases, the other option comes into play: nuclear power generation. By using nuclear power generators it is possible to generate a large amount of power which can continuously feed all the electric systems. The concept of 'space tug' has been chosen as winning candidate for future missions entailing a nuclear power generation system. Here, two the most challenging problems are heat dissipation and radiation shielding (in case of manned missions) and both of them have been addressed and deeply examined along HIPER project. With a constant effort devoted to their further development, space nuclear generators are expected to be fully mature in about 10 years, then making possible a whole new set of space missions (such as the ones requiring a large amount of power very far from the sun).

Main dissemination activities

Besides presentation of scientific outcomes in relevant scientific congresses, conferences and symposia, resulting in 24 publications that will be presented later, HIPER project has been publicised with the public website and in seven conferences/meetings, part of which also open to general public. The website has been publicised with links via the professional social network 'Linkedin' and the general public social network 'Facebook', by using either personal pages and the ones of the coordinator, Alta SpA.

Interdependency with other technology areas

Along the three years of HIPER program, several interdependencies of space related activities with other technology areas have been found. Some relevant example is mentioned hereafter:

1. Electronics: a technology area closely related to spacecraft systems, especially for what concerns the thruster power control unit and, in case of thruster arrays, the switch control unit.
2. Nuclear technology: nuclear power generation was one of the main topics of the program. The study was principally aimed at assessing its capabilities and at developing a new concept of reactor that can be suitable for long range space missions. Of course there are plenty of specific issues connected to such technology which would require a dedicated effort to be tackled.
3. Solar cell technology: solar power was the other source of electric energy considered in HIPER. Studying solar power generation means studying the architecture of solar cells (semiconductor materials) as well as their electric connections and the deployment mechanism of a full SA.
4. Ceramic material technology: in order to manufacture the insulator for the Hall effect thruster a special study has been carried out on the most suitable ceramic materials, investigating their properties and their machinability.

Potential benefits to terrestrial needs

In spite of the nature of HIPER project, which is almost totally devoted to space related issues, some potential benefits to terrestrial needs deriving from this activity can be clearly indentified.

NEO mining:

The international Space community is more and more recognising that a space exploration programme just based on material transported from Earth would be neither affordable nor sustainable. The cost to extract everything from the Earth's gravity well exceeds the world economic capability. Therefore, after an initial phase to build up space infrastructures with material sent up from Earth, it is mandatory to start using Space resources. Besides, there may be a future market for asteroid-derived material (such as nickel-iron grains or semi-conductors like silicon and germanium). Due to diminishing terrestrial resources, a future terrestrial market for precious metals (platinum group metals and gold) may exist, requiring delivery of asteroid material to the surface of the Earth.

NEO hazard mitigation:

We know now that there is the theoretical possibility to mitigate this threat by deviating the trajectory of dangerous NEOs. To implement such a capability it is a matter of developing the adequate technologies. The possibly devastating hazard posed to Earth if hit by a high-energy asteroid or comet is now well recognised by scientists and policy makers. Missions based on high power EP turned out to be an effective option to hook a dangerous near-Earth object and slightly deviates its orbit in order to avoid any risk of collision with our planet.

Nuclear power generators:

Space nuclear generators and space nuclear EP systems are not items that can be developed and implemented within the budgetary frame of ESA. The required investment is so huge that only the whole of Europe can afford it. Very fortunately, such an investment would be a real one in favour of generations to come. Once compact, efficient, safe nuclear generators will be available, deployed in space and tested in very harsh operating conditions, the same generators would be available for Earth applications. Their availability will allow having compact, safe, clean and competitive power plants nearby the communities using their generated energy. Just for this reason a number of positive fallouts may be achieved. By dwarfing to a minimum level the use of lengthy high voltage electrical lines: investments and running costs of power distribution would be radically decreased; heavy energy losses through the grid would be minimised; esthetical and physical countryside pollution nearly totally eliminated, as well as the need for high power high voltage gigantic transformers.

Strategic plans and follow on activities

HIPER represented the first joint effort of a large group of specialised partners aimed at defining the future development of electric space propulsion. Plenty of work has been done in the past three years, plenty of work is still there to be completed in the next years to come. HIPER experience allowed us to set the best direction where to move in the process of making high power EP a winning option for future space missions. However, every single topic can be (and has to be) further expanded, investigated, verified through experimental campaigns, in order to realise items that have the necessary technology readiness level (TRL) to fly onboard a real spacecraft.

The nuclear power generator concept devised under HIPER project has to be physically realised in practice and to be extensively tested, before reaching a level of maturity that can make it ready for a mission. In parallel, a long work of persuasion has to be carried out to underline the intriguing new opportunities offered by a nuclear power generator and to explain that there are almost no risks connected to the launch of such a payload into space (because the reactor would be activated only when in orbit and before the activation it does not emit radiations). This is more politics than engineering, but it is vital as well for the future development of this technology.

The newly conceived solar power generator, which employs solar concentrators (Fresnel lenses) to enhance the flux impinging on a single cell, is already at an advanced development status. A prototype of a complete SA has been realised in laboratory, using a special inflatable structure to support the solar cells and the lenses. The unfolding procedure of the array has been simulated with success, using an engineering model of the solar panel where dummy cells have been integrated. The next step is to move on towards the space qualification of the whole assembly.

MPDTs are probably the devices that are more distant, in terms of development, from the TRL necessary to fly. They require a huge amount of power (although this is not a drawback, because they are supposed to work effectively only at a very high power level and, in theory, to top all the other devices in that power range) and their efficiency is still too low to be competitive with an array of Hall thrusters or GIEs. Understanding the dynamic of the plasma inside such thrusters is the key to shed light on their behaviour and to improve their performance.

GIEs are already commonly used in space and the next steps in their future development concern the realisation and test of the concept of DS3G thruster, theoretically investigated in the HIPER framework. Such configuration is expected to be most beneficial for high power operation of this class of thrusters, increasing both performance and lifetime.

Hall thruster is probably the most mature technology when 'high power operation' comes into play. They proved to be easily scalable to higher power levels than usual (where for 'usual' we mean around 1 to 2 kW) and a full model of a 20 kW Hall thruster has been manufactured, assembled and fired during HIPER program. Results were impressive, with the thruster that immediately operated with a pretty high efficiency (around 60 %) and notable performance in terms of thrust and specific impulse (more than 1 N, around 2 500 s). The thruster is not yet space-qualified and other tests have to be carried out to reach this objective (such as endurance tests to check that the performance do not degenerate after a prolonged firing time).

HIPER public website can be found at

Coordinator's details are the following:

Alta SpA, Via Alessandro Gherardesca 5, 56121 Pisa, Italy

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