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SPAce exploration Research for Throttleable Advanced eNgine

Final Report Summary - SPARTAN (SPAce exploration Research for Throttleable Advanced eNgine)

Executive Summary:
The European vision for space exploration is to explore space with robots and humans. The interest of the Member States in Lunar and Mars exploration has been stressed many times, and one common mission goal that has emerged is enabling returning samples from both within the next decades. To accomplish the above goal a throttling and controllable propulsion system as well as an on ground low cost demonstration capability are needed. These objectives were adopted as main objective of the SPARTAN project.
The SPARTAN project team that includes some of the European top level industries and Academia in the domain has developed a throttling device and a throatable engine capable to use hydrogen peroxide. Thus enabling the development of a higher performances engine able of achieving the above mission lifetime performances. The propulsion system has been designed and developed accordingly and the materials were selected to be as much as possible class “A” compatible with the oxidiser. A tank was specifically developed and its bladder material tested at both specimen and tank level.
The Mission Statement on which the project key requirements were obtained is driven by the Mars mission, requiring soft landing that implies a vertical velocity damping from 30 m/s to 2 m/s. The following mission profile has been derived with also the constrains imposed on the test by the emergency parachute behaviour:
• Release altitude : 170 m; Free fall distance: 50 m (target speed at ignition is 30 m/s);
• Engine ignition altitude: 120 m.
To comply with the above, the SPARTAN Lander Performance Requirements were specified as:
• Impact velocity nominal: 2m/s; Impact velocity recovery: 6.5 m/s Vertical position accuracy: 0.2 m;
• Vertical velocity accuracy: 0.1 m/s.
SPARTAN lander is composed by a primary structure and a landing gear system, housing the following equipment:
• Avionics, including on-board computer, batteries, high precision Laser Altimeter and Attitude Control Unit (Inertia + GPS) and SW, processing algorithms for GNC and managing the sensors (including the ones monitoring the combustion chamber pressure).
• Propulsion System, including one oxidiser tank, two pressure vessels for gas, fill and drain valves, safety and isolation valves, pressure gauges and throttling devices one for each of the four engines.

The Throttling Devices was designed using the cavitating venturi concept that enabled a constant inlet pressure (60 bar) and high efficiency during throttling. Thus the Hybrid Engine (1800N at full throttle) is capable of being throttled down to 180N. It uses HTPB as fuel and H2O2 87.5% as oxidiser, to decompose the hydrogen peroxide before the injection it includes a catalyst, provided by SAAB. The engine design optimisation was achieved by properly set and test the swirling injectors, optimising the axial and the tangential flux of the gaseous oxidiser.
An intensive Development Test Campaign at both equipment and system level has been defined and implemented to verify and validate the performances before performing the final helicopter test.
The throttling device has been successfully tested in a lab environment with water first and then with H2O2, meeting the first important project objective. Afterwards the device was coupled with the engine to verify full engine throttling performance.
The Throatable Engine was tested in several configurations: subscale model at the UNIPD lab, to validate the advanced coding, thus enabling the detailed chamber design; then in the SAAB facility it was tested in Lab model and flight model configuration. The results showed the engine meeting the second important project objective.
In parallel the GNC was validated using a classic stepwise approach: SW in the loop, process in the loop and HW in the loop. This last was performed with the support of a quadricopter first and of a robotic arm after that. Also the delay chain of the recovery system has been characterised by test using an actuator developed for this purpose.
To verify damping performances the Lander system design has been verified through a set of landing gears drop test first and then with a subscale drop test from both a drop tower and an helicopter validating the computed aerodatabase. The wake effect was characterised the by a wake test and a dummy model stability test under the helicopter was performed. Finally the Lander has been tested end-to-end verifying the avionics capability of actuating and controlling the whole system. One restricted test, with the lander attached to a 6DOF load cell, and one quasi-unrestricted test, aimed at lifting-off and hovering the lander, have been carried out to give the green light to the final flight demonstration test.
In this way the partners have acquired the needed know-how for supporting MREP studies and development, being able of deigning a precise landing system with the required system performances. Hydrogen peroxide system and components design and development (useful due to the REACH possible prioritisation of hydrazine) is another important result that will be exploited by the project team.

Project Context and Objectives:
According to the International Space Exploration Coordination Group (ISECG), the ESA /European major objectives and the architecture for exploration are:
• Support European exploration interests, addressing the implementation of European lunar exploration objectives as well as foster technological innovation and Mars-forward preparation.
• Enhance European autonomy.
• Foster stakeholder engagement: create opportunities for international cooperation and broad stakeholder engagement.
• Ensure programmatic coherence: build on European heritage; enable synergies with other ESA space programs and support European coordination towards a targeted role in a global space exploration architecture.
To ensure these high level objectives, it is necessary to demonstrate, key capabilities, solving critical technologies and operational aspects for soft precision landing
SPARTAN project endorsed this strategy developing some key elements aiming at the demonstration of the capability to perform a soft landing, leaving the development of precision landing technologies to a second independent development.
In this perspective SPARTAN research aims are focused at developing a throatable propulsion technology, which is mandatorily needed for any planetary soft and precision landing. This development relies on the hybrid engine technology, exploiting its capability of being throttled within the range of needed thrust range and with appealing specific impulse performances.
The above is translated into the following three project development objectives:
• The engine design, specific for throttling functionality;
• The oxidizer throttling device development;
• The design of the landing case: test bench and testing procedures.
The development was supported by establishing an advanced coding, enabling the definition of the fuel and the throttling behaviour of the core of the project, the hybrid engine.
The first objective has been achieved by realising a 1800N hybrid engine capable of being throttled down to 180N, which uses hydrogen peroxide as oxidiser. This engine has been tested with a manually operated valve with a subscale model at the UNIPD and in a full scale lab model at SAAB facilities. Once the design was optimised, the flight like engine was then tested using the throttling device being developed by Moog Bradford.
The Hybrid propulsion technology was selected, because to its peculiarities (safety, minimum environmental impact (green propellants), low life cycle costs, responsiveness, competitive performances, increased reliability, soft ignition and shutdown) as a fundamental base for the proposed development, since the project team had the knowledge to state that it is capable to provide the necessary engine throttling and shut off capabilities.
The second objective has been achieved through the design of a cavitating venturi valve, providing the hydraulic decoupling. The valve, for safety reasons, has been first tested with water using a test bench specifically developed, and then verified with hydrogen peroxide 70%, results showed a very similar behaviour. Then the valve was coupled with the flight-like engine and the throttling behaviour of the assembly (engine and valve) was verified.
The third objective was the most complex one to be met, it required a wide effort from both the system and the facility development view point, moreover the definition of the mission statement for testing injected unexpected troubles in the development path so that specific wake and stability tests under the helicopter were needed to remove the uncertainties and ensure success.
The system development did also include the parallel development and design integration of the GNC, the propulsion system, the recovery system and the structure, including the landing gears.
The landing test requirements were derived from the soft and precision landing requirements on the mars planet surface (i.e. Exomars and Mars Sample Return missions), adapted to the earth environment:

Mars Environment - Landing Requirements Earth Environment – Landing test requirements
Vertical impact velocity 2 m/s @ 0.5 m from ground 2 m/s @ 0.5 m from ground
Lateral impact velocity 0 m/s Not Controlled
Gravity 3.69 m/s^2 (0.376 g) 9.81 m/s^2 (1 g)
Pressure 600-1000 Pa (about 0.1 atm) 101325 Pa (1 atm)
Total thrust 2500*4 N 1265*4 N
Throttling ratio 1:10 1:10
Total Impulse 100 kNs 12.3 kNs
Spacecraft mass 600 to 1000 Kg 200 Kg

The above preliminary requirements were translated into the following mission statement, driving the system development as follow:
• Release altitude : 170 m
• Free fall distance: 50 m (target speed at ignition is 30 m/s)
• Engine ignition altitude: 120 m
• Impact velocity nominal: 2m/s;
• Impact velocity recovery: 6.5 m/s
• Vertical position accuracy: 0.2 m
• Vertical velocity accuracy: 0.1 m/s
• Throttling engine: 1:10
• Max Thrust: 1800 N
• System dry mass: 320 kg
• Oxidiser capacity: 20 kg
• Firing phase duration: 3.5 s

All the project was under severe costs limitation, so low cost approach was an important driver of the development, which has been considered when procuring equipment, without losing the final project objective.
Any of the above subsystem was tested and validated separately.
The end-to-end capabilities of the SPARTAN Lander, which were the final objective of the entire project, have been verified through a series of system level tests:
• A restricted firing test, with the lander attached to a six degree of freedom load cell to record forces and torques. The lander was commanded in an open loop configuration. Data were collected to verify both the overall delay chain and the synchronicity among the four different engines.
• A quasi-unrestricted firing test, which was planned to be performed twice: in open loop and closed loop configuration. This last test is of a great importance, because it allows to verify all the GNC functions, by introducing the sensors in the loop.

Project Results:
1. Executive Summary 1
2. SPARTAN Final Configuration 3
3. Development Tree 5
4. Development Program and Relevant Results 6
5. System Level Testing 17
6. Objectives Achievements 22
7. Conclusions & Lessons Learned 25

1. Executive Summary
The European vision for space exploration is to explore space with robots and humans. The interest of the Member States in Lunar and Mars exploration has been stressed many times and returning sample from both within the next decades was one of the common mission goals that emerged. In parallel extending human presence beyond Low Earth Orbit toward Moon and Mars has been identified to be accomplished in a step-wise approach.
To achieve the above goals a throttling and controllable propulsion system as well as a low cost demonstration capability on ground are needed. These goals were taken over as SPARTAN’s.
The SPARTAN project team that includes some European top level industries and Academia in the field of propulsion and GNC started almost from scratch, since very little existed up to the time of the proposal in EU and has developed first a throttling device and then a throatable engine able to use hydrogen peroxide instead of nitrous oxide as oxidiser for the hybrid engine taking advantage of the involvement of SAAB as subcontractor thus enabling the development a higher performances engine capable of achieving the needed mission lifetime goals.
The propulsion system was designed and developed accordingly and the materials were selected to have as much as possible class “A” compatibility with the oxidiser. A tank was specifically developed and its bladder material tested at both specimen and tank level.
The Mission Statement on which the project requirements were obtained is driven by a Mars mission, requiring soft landing. This implies a vertical velocity damping from 30 m/s to 2 m/s. The mission profile has been defined after a trade-off between the mission requirements and the parachute recovery system performances, identifying the following requirements:
• Release altitude : 170 m
• Free fall distance: 50 m (target speed at ignition is 30 m/s)
• Engine ignition altitude: 120 m
To comply with the above, the SPARTAN Lander Performance Requirements were specified as:
• Impact velocity nominal: 2m/s;
• Impact velocity recovery: 6.5 m/s
• Vertical position accuracy: 0.2 m
• Vertical velocity accuracy: 0.1 m/s
• Throttling engine: 1:10
• Max Thrust: 1800 N
• System dry mass: 320 kg
• Oxidiser capacity: 20 kg
• Firing phase duration: 3.5 s.

The SPARTAN lander is constituted by a primary structure and a landing gear system, supporting the following equipment:
• Avionics, including on-board computer, batteries, high precision Laser Altimeter and Attitude Control Unit (Inertia + GPS)
• Software, processing algorithms for Guidance, Navigation and Control in a closed loop, and managing sensors (including combustion chamber Pressure sensors) and actuators (throttling devices);
• Propulsion System, including one oxidiser tank, two pressure vessels for gas, fill and drain valves, safety and isolation valves, pressure gauges and throttling devices, one for each engine.

The Throttling Devices was designed applying the cavitating venturi concept, because of the major advantages given by hydraulic decoupling: varying the mass flow (600g/s to 60g/s), the downstream pressure does not vary. A constant inlet pressure (60 bar) is needed to maintain high engine efficiency during throttling.
The Hybrid Engine (1800N) is capable of being throttled down to 180N. It uses HTPB (NAMMO proprietary) as fuel and H2O2 87.5% as oxidiser. It has an average specific impulse of 230 s, over the full throttling range. The combustion chamber has a length of 165 mm and a O/F ratio rages from 7.5 to 8.5. It includes a catalyst, provided by SAAB, needed to decompose the hydrogen peroxide before the injection. The engine design optimisation was achieved by properly set and test the swirling injectors, optimising the axial and the tangential flux of the gaseous oxidiser.

SPARTAN project implemented an intensive Development Test Plan at both equipment and system level, verifying the achievement of the established project objectives before performing to the helicopter test.
The throttling device was tested in a lab environment with water first and then with H2O2. The results showed performance compliance, meeting the first important project objective. Afterwards the device was coupled to the engine to verify full engine throttling performance.
The Throatable Engine was tested in several configurations. Using a subscale model at the UNIPD lab, who realised a complex setup, which allowed them to reproduce the H2O2 stoichiometry. This test validated the advanced coding, enabling the detailed chamber design. Then it was tested again at SAAB facility, using a lab model and a flight model. During this second campaign the mass flow was driven by the actual throttling device. The results showed the system meeting the second important project objective.
The GNC was validated using a classic approach: SW in the loop, Process in the loop and HW in the loop. This last was performed with a quadricopter first and a robotic arm after that, during the second tuning loop, before the TR4 campaign.
The delay chain of the recovery system was characterised by test using the actuator developed for this purpose.
The Lander system design was verified through: landing gears drop test, verifying damping performances; a subscale drop test from both a drop tower and an helicopter, validating the computed aerodatabase; a wake test and a dummy model stability test under the helicopter, characterising the wake effect.
Finally the Lander was tested end-to-end to verify the capability of the avionics of actuating the four engines simultaneously, synchronised and with a given profile. The test was carried out successfully with the lander attached to a 6DOF load cell.
The last step includes the lift-off and hover the lander, laterally restricted, to verify its closed loop behaviour, managing the vehicle dynamics, having the sensors in the loop. This test demonstrate the last objective of the project, giving green light to the final flight demonstration test.
Having achieved the above the partners acquired the know-how of supporting MREP studies and development, increasing the system capabilities towards a precision landing as well. Hydrogen peroxide system and components design and development (useful due to the REACH possible prioritisation of hydrazine) is another important asset that will be exploited.

2. SPARTAN Final Configuration
The system is composed by the following sub-systems:
• Structure
• Legs and gears
• Propulsion
• Guidance, Navigation and Control
• Recovery System
• Image recovery and Telemetry.

The System performance characteristics are:
Touch down velocity: 2 m/s
Vehicle velocity at engine ignition: 30 m/s
Throttling engine: 1:10
Thrust: 1800 N
Fuel: HTPB
Oxidizer: H2O2 87.5%
Maximum Mass flow rate: 0.6 kg/s
System dry mass: 320 kg
Oxidiser capacity: 20 kg
Firing phase duration: 3.5 s.

The Propulsion System includes:
1 Oxidiser tank (bladder type)
2 Pressure vessels (COPV)
1 High Pressure Transducer
2 Low Pressure Transducers
1 Fill and vent valve
1 Fill and drain valve
1 Pressure regulator
1 Pressure relieve valve
1 Safety relieve valve
1 Oxidiser Tank vent valve
1 Rupture Disk
1 Oxidiser relieve valve
1 isolation valve
4 Flow Control valve (throttling 1:10)
4 Hybrid engines (HTPB fuel, including Combustion Chamber Pressure Sensor).

The Engine includes:
Catalyst pack for H2O2 dissociation
Whirl Injector
Motor case and external interfaces
Fuel cast

The avionics system includes:
Guidance Navigation & Control electronics ad algorithm
1 IG500N
Laser range
On Board Computer
Power Distribution Unit
Flight management

The Ballistic Recovery System includes a GRS-450 hard pack that is a ballistic parachute system from Galaxy High Technology System designed for ultra-light aircraft with a maximum take-off weight of up to 475kg.
The currently selected parachute system has the following features:
• Maximum payload: 450 kg
• Weight: 13.3 kg
• Able to be deployed up to 160 km/h
• Speed of descent: 6.5 m/s (450 kg)
• Ejection time: 0.8 s (rocket motor).

The structure and the legs are of a tubular type, made out of Aluminium.
The dampers are designed and manufactured by QUIRI according to the following reference requirements
• Number of dampers : 4
• Maximal mass to damp : 350 kg
• Maximal vertical impact velocity : 6.5 m/s (impact velocity in case of parachute actuation)
• Maximal damper stroke : 270 mm.

The Oxidiser Tank
The oxidizer tank is being developed and manufactured by EPE Italiana S.R.L. following an innovative approach to reduce developing and manufacturing cost, maximizing the use of existing technology and off the shelf products already available in other engineering fields.
The oxidizer tank technology is being borrowed from hydraulic bladder accumulators, commonly used in hydraulic applications, to damp the pressure fluctuation thank to a reservoir of a pressurized gas (usually nitrogen) contained inside the bladder.
Type: Bladder accumulator with transfer valve
MEOP: 75 bar
Test pressure: 115 bar
Burst pressure: 220 bar
Nom. Capacity: 20 dm3
Shell material: Inox AISI 316L
Bladder material: FKM fluoroelastomer.

3. Development Tree
SPARTAN has been ruled through the following Development Tree, which explains into details how the project has been developed in order to achieve the objectives:
1. Mission statement
Performance requirements
Recovery System
2. Fuel/Oxidiser baseline
Throttling device
Regression rate sensor
Subscale testing with gas generator (purpose development)
Catalyst unit

3. Flight Dynamic Simulator
Architecture definition
Algo definition
Wake & stability
Actuator characteristics
Sensor characteristics
Code generation

4. Avionic system: Sensors Selections and characterisation

5. Structure and landing gears

6. Oxidiser Feed system

7. Oxidiser tank

8. Ballistic Recovery System

9. System Level Testing
Wake & Stability
Landing gear Impact 1
Landing gear Impact 2
Drop tower (subscale model)
Stability under Wake (full scale model)
TR3A (only water)
TR3B (firing)
TR4A  Open Loop Dynamic System Firing Test
TR4B  Closed Loop Dynamic System Firing Test
Demonstration Flight Test

4. Development Program and Relevant Results
1. Mission statement
The mission statement has been identified by trading off two conflicting operating requirements:
The originally proposed mission statement, which is related to the final goal, and the recovery system behaviour, which was put on board to enable the safe recover any anomaly that could occur during the flight.
The basic mission statement was: reach 30 m/s of vertical velocity, than damp the velocity down to 2 m/s at 0.5 m from ground. To achieve 30 m/s 50 m of free fall are considered enough, while damping to 2 m/s is required at about 100 m, but this last length was not enough to guarantee the full deployment and operations of the parachute system, due to the actuation chain delay.
Therefore, after the actuator characterization test, was modified as follows:
• Release altitude q0: 170 m
• Free fall distance: ~ 50 m (target speed at ignition is ~ 30 m/s)
• Engine ignition altitude q1: 120 m

2. Fuel/Oxidiser baseline
At the beginning of the project there was a debate within the Consortium, to decide whether to pursue with the proposed combination of fuel/oxidizer (paraffin / Nitric oxide) or to move to another selection: HTPB and hydrogen peroxide 87.5%
The proposal came from NAMMO and it has been justified by their wide experience in managing the HTPB and their interest in starting dealing with the hydrogen peroxide for both applications: hybrid rockets and monopropellant thrusters. NAMMO also supported the proposal with expected better engine performances.
With this purpose NAMO established a relationship with SAAB Under water systems, to spin in their experience in managing the hydrogen peroxide.
A trade-off has been carried out, according to the following parameters:
• L/D engine ratio
• Combustion efficiency
• Material compatibility over time
• Oxidizer performance degradation
• System complexity
• Storage system availability/maturity
• Reliability
• Safety on ground
• Costs
The material lifetime compatibility was one of the biggest concern that driven the trade-off. The investigation verified that the hydrogen peroxide storage lifetime is compatible with the reference mission profile: mars landing and return. Having that in mind and the better performances and throttling capability of the engine in case the hydrogen peroxide would have been selected, the Consortium in its whole decide to move to the HTPB/H2O2 combination.
This choice influenced the development of the throttling device, the throttling demonstration approach and the oxidizer feed system as well, according to the following item list.

Throttling device
The valve is based on cavitating pintle technology, designed to ensure cavitation over the whole mass flow throttling range (0.1 ÷ 0.6 kg/s). this guarantee a stable engine inlet pressure condition over the complete mass flow range: from full throttle down to minimum one. Therefore the Mass flow is independent of downstream conditions of the valve, i.e.:
• Mass flow is unaffected by chamber pressure changes during transients from ambient to operating pressure.
• Mass flow is unaffected by combustion oscillations during steady motor operation.
The pintle is dynamically positioned by means of an electric servo-motor EC22 which is supplied and controlled by a separate Maxon EPOS2 24/2 (p/n 380264) digital positioning controller. There is one dedicated controller for each valve, which receives the main control input directly from OBC.
The EPOS2 controller is not integrated within valve body. The four controllers shall be enclosed in a dedicated case.
The valve design and performances have been verified through the test sequence 1 &3 with water and the test sequence 2 with hydrogen peroxide.
The primary objective of the test sequence 1 are to quantify the mass flow characteristic of the valve and to determine the relationship between pressure forces acting on the pintle as a function of its stroke. The former is used to establish empirical parameters required for valve modelling in support of the flight design and the latter is required for sizing of the actuator.
Secondary objective is the determination of the leak-tightness of the valve.
The test item consists of breadboard flow control valve. Two pintle and two nozzles were defined.
The test sequence 3 was performed on the breadboard model 2 flow control valve.
The breadboard model 2 is a modification of the existing breadboard model 1 hardware. This model incorporates a DC motor to drive the pintle. The primary objective of this test sequence was to verify the functioning of the actuator and to obtain first estimates of the attainable response times of the valve/actuator combination.

The primary objective of the test sequence 2 was to characterize the flow rate through the valve as a function of pintle position and back pressure when supplied with H2O2.
Performance is evaluated by direct comparison with results obtained on the BBM2 (test sequence 3)

From the test results it can be concluded that:
• The valve in its present configuration is externally leak-tight to the point that there is no visible liquid Leakage.
• A pintle with a parabolic contour results in a linear relationship between mass flow and pintle position when the valve is operating in cavitating mode.
• Once closed, the force of the valve spring is sufficient to maintain the closed configuration.
• The actuator therefore does not need to be active to ensure continued shut-off.
• Flow control accuracy of 0.9 % FS (1confidence interval) was demonstrated.
• Hysteresis is limited to 2 % FS and this corresponds to worst-case expectation.
• Step changes of 20 % of the full-scale control range was realized within 0.1 s and demonstrated.
• No overshoot was detected at any time when subjecting the valve to step demand inputs.
• Flow control and detection resolution was better than 0.002 g/s.
• Measured water mass flow rate to hydrogen peroxide mass flow rate conversion agrees to within 1.0 % of theoretical straight-line predictions based on the linear fit.
• No adverse effects of the hydrogen peroxide on the flow control valve could be determined.
Subscale Engine testing with gas generator (purpose development)
The engine design and performances have been validated through three different models:
• 1 subscale model test
• 1 Full Scale Lab model test (heavy weigh) – TR1
• 1 Full Scale Flight weigh model test. This last has been tested twice: with a manually operated throttling valve and the flight design throttling valve – TR2A &B

The main aim of subscale model is to conduct experimental tests on a lab scale motor of the SPARTAN propulsion system, to acquire data for the CDF code, developed by UNIPD. The requirements have been identified on the basis of CFD performance, i.e. from the output data of the code, which have to been validated.

From the test results it can be concluded that:
The subscale motor simulate the oxidizer using a gas generator, which provides a hot mixture flow rate to the injector inlet, with composition as similar as possible to decomposed H2O2 at 650°C. The gas generator comprehends a mixing chamber, in which oxygen and liquid water are injected and heated up by a O2/HDPE hybrid torch. The mixture is then injected in the main combustion chamber by means of a vortex injector.
This subscale motor has provided experimental data, which validated the numerical CFD code.
Finally, experiments have been performed on the complete engine for the 40% throttling level to verify the HTPB behaviour during full duration firing tests: the fuel demonstrated ignition transient as it was expected, and a stable combustion with the vortex injected oxidizer mixture.

Full Scale Engine testing
The TR1 had the scope of preliminary verifying the throttling capability of the engine, validating the design.
The goal of TR2 was to demonstrate the flight weight rocket motor design based on results found during TR1. There will be two different phases:
Phase A utilized SAAB’s hydrogen peroxide (H2O2), run-tank, valves, and recording equipment as performed in previous programs in cooperation with Nammo.
Phase B utilized the Moog Bradford developed pintle throttle valve instead of the SAAB throttle valve. Other than this, the same equipment was used for transferring H2O2 to the engine and for measuring the various rocket parameters. Phase B was carried out some months after that phase A has been completed understanding the start-up behaviour in terms of both the delay itself and the spread between the different engines.

From the test results it can be concluded that:
During firing tests the engine performed as expected over the whole throttle range.
The TR1 sequence validated the engine design and the catalyst pack one.
Some ignition delay has been detected during the TR2B, which is addressed to the test set-up rather than to the engine itself. In detail:
• 15 engines has been tested within the SPARTAN TR2 campaign. 10 engines has been tested with SAAB throttle valve while the remaining 5 has been tested with Moog Bradford battleship throttle valve.
• Four different fuel geometries have been tested in an attempt to reduce time to ignition. The conclusion in respect to best geometry is non-conclusive.
• TR2B encountered to high reduction in tank pressure during full throttle mode due to a too high flow restriction in the gas pressurization system (200bar system adjusted to 70bar for the Moog Bradford throttle valve). This resulted in less oxidizer mass flow being injected. As a consequence this also led to increased time to ignition. TR2A did not have this due to that a high capacity flow line was used, however at lower pressure, but adequate for the SAAB throttle valve.
• Main conclusion: The ignition timing seems to be too unreliable using only a catalyst unit SPARTAN needed a heater or igniter system.
• Testing of an ignition and/or a heater unit occurred during TR3.

Regression rate sensor
The project requested to develop a diagnostic equipment for real-time local regression rate measurement, through:
• Selection of possible systems from the market or design of a specific apparatus,
• Implementation and test of the method in a lab-scale model engine,
• Define the implementation methodology into a real fuel grain.
Existing options
• Commercial system MIRRAS from ORBITEC: not available for purchase outside USA territory.
• Ultrasound/microwave technique: expensive apparatus and long tuning process required.
• X-ray technique: expensive for implementation and safety requirements.
A dedicated wire-cut regression rate measurement has been finally developed (patent pending)
• The sensor is embedded in the fuel grain:
o As the fuel is consumed, the wires are broken by the flame,
o The resistance of the circuit is changed,
o Measured voltage across R changes.
• The electrical resistance will change during combustion.
• Five different generations of sensors have been developed.
• Two different configurations are available depending on fuel grain loading.
• Minimal connections are requested for easier operations.

The sensor has been validated in the PoliMI lab facilities first, than on the subscale model. It was finally used embedded into the full scale engine at the SAAB facility, providing very useful results:
• Grains have been instrumented in three locations
• The resulting signals were clear
• In the reported example, the grain is consumed faster at the head end of the motor
• Information on local fuel grain consumption could be obtained

Catalyst unit
The selected 87.5% hydrogen peroxide purity grade has excellent decomposition capability when using silver screen catalyst material.
The catalytic decomposition of hydrogen peroxide is generally simple:
87.5 weight % of 1 mol H2O2(l) gives 0.788 mol H2O2(l) + 0.212 mol H2O(l)
0.788H2O2(l) + 0.212H2O(l) 1H2O(g) + 0.394O2(g) + heat
Decomposition temperature is typically 660C for the concentration of interest. The catalyst pack actual size has been based on the SAAB torpedo experience:
Maximum oxidizer mass flow 87.5 % H2O2 0.7 kg/s
Maximum pressure drop over catalyst pack < 8.5 bar
Length of catalyst pack 50 mm ± 10 mm
Diameter of catalyst pack (internal) ± 90 mm
Catalyst material Silver
Front distribution plate Radial holes
Structural material Nimonic 75

The ignition delay that has been verified during the TR2-B engine test sequence, led to introduce either heaters on the catalyst pack or igniters into the fuel grain.
A heater pack has been tested, with limited success; therefore igniters have been adopted and installed into the fuel grain for the TR4 system test sequence.
Their characteristics are:
0.8-1 A; +12VDVC; 4 ms

3. Flight Dynamic Simulator
The FDS is not only the development base for embedded software, but is also the main analysis tool employed for flight dynamics investigations, as it reproduces not only the control segment but also the guidance and navigation aspects and manages environmental variables as well.
The navigation is composed of two main blocks. The position and velocity block is in charge to estimate the position and velocity of the SPARTAN experiment, and the estimation is performed using a discrete Kalman filter. The attitude block estimates the attitude, but this block contains only a bus selector because the attitude is computed directly using the gyroscope measurements.
FSW development took advantage from auto-coding technology and is directly derived from Spartan FDS. The on-board software architecture is composed of three elements:
1. The sensors’ interface;
2. The GNC onboard software;
3. The actuators’ interface.
The main element is the GNC onboard software, which implements two modules (see Figure 4):
The ALT module runs at 100Hz frequency; pre-processes the measurements provided by the Laser Rangefinder;
The FSW module runs at 10Hz frequency; implements the main functions GNC functions:
The Guidance function;
The Navigation function;
The Control function;
The Flight Management function;
The Body function.

The SPARTAN Algorithm, which identifies the design of the Flight Dynamic Simulator, is constituted of the following parts:
Flow Control valves
Fuel Grain
Aerodatabase. Spartan aerodynamics was defined, by means of CFD analysis and validated by means of a drop tower test
Wake & stability
Avionics, including sensors
Inertial sensors
Temperature Sensors
Data processing
On Board Computer

4. Avionic system: Sensors Selections and characterisation
The following major requirements drove the GNC design and components selection:

Navigation accuracy:
The Navigation System provides the kinematic state of the vehicle with the frequency of 10 Hz and the following accuracies (RMS):
• Position (xy-axis): ≤ 3 m
• Position (z-axis): ≤ 0.02 m
• Velocity (xy-axis): 0.5 m/s
• Velocity (z-axis): 0.1 m/s
• Attitude: 0.5 deg
• Angular velocity: 0.1 deg/s

All values are considered at 1-sigma.

Altitude measurement:
SPARTAN is equipped with an altimeter capable to provide a measurement accuracy of 0.03÷0.05 m over a minimum distance range of 170 m.
• The Guidance System shall provide, with a frequency of 10 Hz:
• the ignition command (only once, no re-ignition)
• the following commands for each of the 4 engines:
o throttle percentage.
o max rate = 200%/s
Temperature: -5°C till + 34°C
Humidity: 24% till 100%
Dew point: -26°C till ÷ 21°C
Pressure: 978 till 1037 hPa
Dust and cement landing site
In light of the above, the Avionics/GNC architecture is organized as follow:
The main constitutive components are:
(i) Navigation and Sensing:
o Automatic safe and arm switches (providing a status/mode feedback)
o Diagnostic sensors (pressure, temperature and voltage sensors)
o Navigation sensors (IMU, AHRS, GPS, Altimeter)
(ii) Guidance, data storage and transmission:
o On Board Computer (OBC)
o Data acquisition and storage units (DAU, MMU)
o Wireless modem (i.e. telemetry)
(iii) Power and distribution:
o Batteries
o Power distribution (PDU)
(iv) External interfaces
o Manual control panel (2 power switches plus 2 mode/program selectors)

During the development it has been noticed that the LIDAR could have been affected by the engine pollution.
An analysis has been carried out considering the engine exhausts concentration defined by the advanced code developed by UNIPD and it turned out that those elements should not have jeopardized the LIDAR performances.
Nevertheless a test has been planned during the TR2, at engine level. The objective of the test was to understand the behaviour of the laser altimeter measurements with the interaction of the plume of the motor that will provide the thrust at the SPARTAN lander experiment. The idea was to characterize the possible variation of laser range when the beam of the laser is passing through the plume of the motor, varying the angle between the plume and laser beam. This test has been performed at SAAB building during the firing test of NAMMO with a LIDAR installed perpendicular and parallel to the engine plume direction.
After performed the firing test, there are two sets of laser range measurements: the measurements without the plume, and the corresponding with the plume of the motor. The difference between the two sets of data represents the error introduced by the plume in the laser range measurements. The interaction between the plume and the LIDAR has been experience during the early monopropellant phase only, therefore not influencing the operating phase of the landing.

The IG500N and the LIDAR have been characterized with devoted test campaign on ground.
The results so far indicate that the added value of the IG-500N is its ability to integrate different sensor measurements and output a coherent navigation solution.
It was seen that in outdoor environments, when GPS has full visibility, the GPS performance is better than the performance of the IG-500N navigation, but that navigation has more consistent accuracies throughout the tests.
It appeared clear that the IG-500-N is not able to fulfil most of the navigation requirement and the use of the Laser Altimeter is deemed necessary, which was selected and tested successfully.

The validation at system level has been carried out at both GNC and lander level.
The validation at GNC level has been carried out by means of a quadricopter (UAV), on which sensors, OBC, and relevant SW, have been installed.
The validation at Lander level has been carried out during the TR3-A testing at NAMMO: end to end test up to FCVs, with the tank loaded with water.
The valves have been commanded with a defined low. The mass flow has been recorded to verify proper throttling set.

A quadricopter test flight has been performed by GMV in order to validate the coupling among the LASER altimeter and the attitude determination.
The baseline used to assess the performance of the sensors, of their coupling and of the developed algorithms is the data collected by the Spyro independent avionics, provided by UAVision. The presented data, spread on a time window of 8 minutes and 20 seconds, show very good correlation between the SPARTAN GNC system and the independent avionics.
The comparison between positioning, represented in Latitude, Longitude and Height (LLH), computed by the INS and the UAV avionics are presented in Figure 8 2. As it can be easily seen, there is a small offset between the three components of the two data sets. The mean error is, respectively for latitude and longitude, 1.26e(-4) and 4.18e(-5) degrees. This can be translated to a mean error of approximately 14 meters south wise and 4.5 meters east wise. In terms of altitude, the UAV avionics reports a height that is, on average, 4.22 meters higher.
The relative position on the ground, computed in post-processing using the data presented in the previous section, from the INS is compared with the relative position provided by the quadcopter avionics.
The navigation results use the laser altimeter measurements to compute the altitude, unless at least one of these two events occurs:
• At least one of the LR measurements present in the buffer is out-of-range (reads 4095 decimeters);
• The bank angle with respect to the vertical, measured by the INS, is higher than 30 degrees.
The first event was recorded twice. This can be caused either by the laser beam being reflected by a reflective surface (e.g. polished metal, water puddle, etc.), or, by the opposite, by being absorbed by an absorbing surface (e.g. grass, tree leaves, among others).
The second event occurs 4 times.
If any of the previous events is observed, the navigation algorithm replaces the method of computing the altitude using range measurements by the INS internal navigation altitude. The jumps in navigation altitude brought by these changes can be observed at the points in which the navigation altitude equals the IG-500N altitude.
In any case, apart from the discussed events, the navigation altitude, computed with LR measurements, is very similar to the altitude provided by the Spyro’s avionics and the barometric altitude

5. Structure and landing gears
SPARTAN structure is constituted of a frame of Aluminium tubular trusses.
Bracket and supports have been designed according to the worst load case, when possible: Impact velocity = 6.5 m/s
On top of structure a hooking ring is provided to allow lifting of SPARTAN by means of helicopter
The geometry is designed to:
• Absorb the forces uniformly
• Allow easy installation of different components
For best combination of weight reduction and strength purposes, all parts have been manufacture of Aluminium 6061 T6.
The structure have been designed and validated by FEM analysis.
For structural validation, the following cases are taken into account:
• Hoisting / transportation phase: the FEM is fixed at the top and the following combined loads are applied on the whole structure:
o 2G on Z axis;
o 1.5G on Y & X axes combined.
• Motor thrust phase: an accelerations 15.63 m/s2 in +Z direction is assumed and applied on the whole structure.
• Landing: two load conditions are considered:
o Static load: the case marks out the boundaries of the low position and the high position under a 5G solicitation with the I/F feet / ground blocked along out of plane direction and radial direction.
o Dynamic load: a velocity of 2.5 m/s under 1G is assumed.
As a result, the primary and the secondary structure are validated by the analyses.

The base structure is covered by flat panels, In order to obtain aerodynamically stable shape and surfaces. The main purpose is to ensure flight stability during the free fall phase. Such stability is expected to be obtainable with no need of active stabilization systems (e.g. RCS) with AOA within a range of α = 0°÷30°

The lander is equipped with four legs and each leg is equipped with a damper tube.
The dampers have been specified by STUDILE, designed and manufactured by QUIRI, according to the following major requirements:
Maximal mass to damp : 350 kg
Maximal vertical impact velocity : 6.5 m/s
Maximal damper stroke : 270 mm

The design damping to elongation characteristics that has been defined is a nonlinear one, ranging from 200 to 1800 Ns2/m2 with a half parabolic shape. The dampers behaviour has been verified by means of three drop test campaign and the performances correlated with the test design.

6. Oxidiser Feed system
The functional configuration consists of a pressurization section and an oxidizer feed section, the latter subdivided into four identical branches, with each branch feeding a single engine.
The pressurization section consists of a high-pressure pressurant tank, which stores gaseous nitrogen at a pressure of up to 200 bar. The pressurant is fed through a pressure regulator to the propellant tank to maintain propellant pressure at a constant 65 bar. The pressurant tank can be filled vented (or vacuum can be made as well) through a fill and vent valve. Pressure is monitored both upstream and downstream of the pressure regulator. In case of regulator failure with a filled pressurant tank, a relief valve is installed downstream of the pressure regulator. A further solenoid latching valve is provided in order to allow safe, remotely operated, de-pressurization of the system.
The oxidizer feed section consists of an oxidizer tank, which stores hydrogen peroxide at a pressure of 65 bar. For safety purposes, the tank is provided with a vent branch that prevent overpressures during both loading (through a ball valve) and flight operations (through a (primary) relief valve and a (secondary) rupture disc). Oxidizer is fed to four engine branches through a common NC solenoid valve, which allows isolation of the propellant tank from the engine branches. Each engine branch contains a flow control valve, implemented by means of a variable orifice cavitating venturi. Each flow control valve doubles as a mass flow meter thanks to its independence of the downstream pressure.

For ease of integration and simplicity, the fluidic system is connected by means of Swagelok fittings to the largest extent possible. Components internal to the oxidizer feed system are interconnected by seamless bright-annealed tubing compliant with ASTM A269.
Dimensions (indicated with nominal diameter x thickness) for the selected tubing are:
• Pressurant side:
o high pressure (200 bar) tubing from pressurant tanks up to pressure regulator: ½” x 0.065”.
o low pressure (65 bar) tubing from pressure regulator up to oxidizer tank: ½” x 0.065”.
• Oxidizer side:
o tubing in the common outlet section from the oxidizer tank up to the Tee from which the two symmetrical rocket engine branches diverge: 1” x 0.120”.
o tubing between the diverging Tee up to rocket motor final branches tees are ¾” x 0.095”.
o terminal branches (including FCVs and rocket motors) are ½” x 0.065”.
• Load/Purge/Vent branches are all ½” x 0.065”.

The analysis of system overall pressure drop has been carried out.
A detailed model of oxidizer feed system have been built by means of EcosimPro. The simulations are devoted to the following analyses:
• Local and overall pressure drop.
• Pressurant and oxidizer blow-down.
• Line priming effects and water hammers
With regards to flow control valves, two feed-system architectures have been initially proposed:
• Single stage flow modulation: composed of a dedicated flow control valve for each motor, which is able to perform full range mass flow modulation with sufficient accuracy;
• Double stage flow modulation: provides one further flow control valve to manage the coarse flow rate modulation (gross thrust changes). This valve is added before the four dedicated FCVs, now dedicated to fine flow rate modulation only (attitude control).

Single stage architecture provides linear relation between thrust modulation and valve Cv variation.
Double stage architecture requires a bigger throttling ability and the relation between thrust modulation and fine control valves Cv variation is highly not linear and asymmetrical. Single Stage was selected.

7. Oxidiser tank
SPARTAN oxidizer tank is based on standard accumulators concept but its special safety and technical requirements involves some deviation from the standard design. Such deviations require to be assessed by testing in order to verify that there are no impacts on the expected performances of the tank. The modifications with respect to COTS design are:
• Removal of the anti-extrusion spring, that will be compensated by reducing the diameter of holes on the anti-extrusion tube and increasing their number (8 rows of 16 bores Ø2 mm). The purpose of the change is to minimize the surface to volume ratio.
• Use of a dedicated FKM compound for the bladder, which will allow major elongation capability and chemical compatibility with 87.5% H2O2.
The main purpose of the test is mainly to assess and validate tank performance with respect to
SPARTAN mission requirements:
• Actual water capacity
• Pressure drop at the outlet

The tank is polar mounted:
• The upper interface (N2 inlet and H2O2 vent) is not constrained in the axial direction.
• The lower interface (H2O2 outlet) is fully constrained.
The tank is compliant with 97/23/CE Pressure Equipment Directives (PED).
The shell is checked for structural integrity by means of radiography.

The following procedure was applied during testing:
- Loading of 20 kg of water;
- Pressurization (static) at 75-80 bar;
- Full opening of the start/stop ball valve;
- Throttling of the FCV to achieve the desired flow rate;
- Closure of the FCV before discharging more that the safe amount of water (to avoid bladder extrusion).
Four tests were performed at several flow rates, as specified in the following table. Water flow rate was dynamically adjusted during each test, thus the values reported correspond to a quasi-static behaviour.
The pressure drop was measured between the gas inlet and water outlet of the Spartan tank. Tests were made at several flow rates, set between 100% and 10% of the nominal mass flow rate (2.4 kg/s of H2O2). The drop was always below 1 bar. The bladder gave an overpressure of maximum 0.5 when it was fully loaded. This overpressure was negligible when more than 10 lt of water were discharged. The maximum amount of water loaded was 19.8 lt, achieved with a loading overpressure of 2.5 bar.

The Bladder material has been tested to verify the compatibility with H2O2 87.5%.
The selected bladder material is FKM (FLUOROCARBON Elastomer). The chemical resistance to H2O2 has been tested by SAAB, performing immersion tests on both unstressed and stressed coupons. SAAB applied international standard as test specification. The coupon has been cut out from a bladder coming from the same flight batch. The results confirm the material is class 1 with respect to H2O2 87.5%. The lubricant Christo Lube has been tested as well by SAAB, giving positive results.
The compatibility test has been repeated at bladder level, showing a slight degradation of the decomposition rate, but still placing the material at a value belonging to a good class 2 material.

8. Ballistic Recovery System
The design of recovery system is performed according the aforementioned architecture and applying to following requirements:
• Minimum load capacity of 350 kg
• Minimum weight of the system
• Safe to open up to 150 km/h
• Lowest possible speed of stable descent
• Minimum possible parachute opening time.

The selected parachute system has the following features:
• Maximum payload: 450 kg
• Weight: 13.3 kg
• Able to be deployed up to 160 km/h
• Speed of descent: 6.5 m/s (450 kg)
• Ejection time: 0.8 s (rocket motor)
• Opening time: 1.8÷2.5 s (no slider); 2.7÷3.1 s (slider)
The BRS actuation takes advantage of pyro-actuators.
All of main system parts underwent through testing in laboratory and real conditions.
Laboratory test were consisted from PA force estimation, firing time analysing and safety conditions needs.
Main aim of the real condition test was to prove function of BRS as a complex system.
Secondary targets were to clarify :
• Covering cap separating
• parachute canopy deploying trajectory
• deployment time sequence
• rocket trajectory and maximal altitude.
Rocket ejection time has further been confirmed by means of a dedicate ground tests.
The test also provided useful design feedback for BRS remote actuation system design improvement.
It also confirmed that BRS is safe and applicable to Lander. Measurement of rocket trajectory and travel distance also confirmed that there’s no risk of collision with helicopter.
Actuation delay has been finally characterized.

5. System Level Testing
i. Wake & Stability
Scope: Verify the aerodynamic behaviour of the environment under the wake, to manage the recovery phase of the mission.
Conclusion: The effect of tail rotor is unknown. The variation of wind (even small velocity like 1ms-1 with altitude unknown). Motion of helicopter to keep position induced disturbances. Probe oscillates during test (pendulum).

The test did not give direct confirmation or disagreement with computational predictions.
There was measured stability of probe during test and was successfully tested equipment – telemetry.

ii. Landing gear Impact: run 1 & run 2
Scope: Verify the landing gears design at increasing impact velocity: up to 6.5 m/s, which is the maximum design impact velocity, derived by the worst case landing impact scenario: recovery system activated

Conclusion: During drop test run1, there was a severe failure of the dampers and legs, which broke. During the second run, after both legs and damper redesign, the test was stopped at 3.5 m/s impact velocity, because of the still high impact loads registered. It was decided to not damage the legs.
The results of this test lead to modify the damper design one more time.

iii. Subscale Model Test: Drop tower and Helicopter
Subscale Model Physical properties
Scale 1:4
Characterize dimension: width [m] 0,51
Mass [kg] 3,733
Moment of Inertia: Ixx [kg*m2] 0,0641
Moment of Inertia: Iyy [kg*m2] 0,0641
Moment of Inertia: Izz [kg*m2] 0,1090

Scope: Validate with an instrumented subscale model, released from a drop tower, the aerodatabase and the full scale model dynamic behaviour determined by CFD

Conclusion: CFD coefficients properly validated

The subscale model was also tested from the helicopter.
In the order to investigate SPARTAN demonstrator behaviour in free-fall phase the model was built up. Model is scaled in ratio 1:4 to original and was already used for laboratory drop tests with specific initial condition.
Geometry of SM, moments of inertia, mass distribution and CoG position is set as close as possible to original. However model is simplified about landing gear and surface details. Its low affecting for overall aerodynamic characteristics was proved by CFD simulations.
Model construction is assembled from plywood pieces where edges are stiffed by glass fibre laminate. Dividing plane in the widest point allows to model be divided into two separate parts.
Drop test altitude: According to SM simulations, laboratory drop tests and LD stability test the drop test altitude is set for 400 m, where the altitude of SM is 320m.

Scope: The main purpose of test is to verify influence of helicopter rotor wake on initial dynamic state of lander before release and on lander stability during free-fall phase.

Conclusion: Due to the helicopter size, which was not a heavy weight one, and the lander size, too small scale for such a test, it didn’t provide reliable information, mainly when the free fall lander stability is concerned. Insofar this issue the small scale generated an unexpected torque due to the friction of the hook release mechanism.
One important outcome from this test is the necessity to use a heavy weight helicopter, enabling a better stability during hovering, before the lander release.

iv. Stability under Wake (full scale model)
A test has been carried out, using the dummy structure without legs, no release.
The scope of this test was the verification of the wake influence on a full scale model under the helicopter.
This test revealed that the selected cable length (80m) was properly set, because no remarkable attitude anomaly was detected during the hovering, which was simulating the pre-release phase.

v. TR3A (only water)
Scope: The water testing has been planned with the lander in flight configuration, but the landing gears. The scope of the test is to verify the proper behaviour of both the propulsion system in quasi end-to-end configuration, in safe mode: engines only are missing.
A second important objective is to verify the safe loading procedure, using distilled water instead of H2O2.
The FCV’s have been commanded by the OBC, verifying a given throttling profile. The test was run twice, because, due to a provisional loading equipment, the IV has been polluted, generating a gross leakage.
The IV has been dismounted and repaired. The second run has been carried out with GNC sensors included: LIDAR has been tested as well.

Conclusion: After executing the TR3-A, run 1, the results were analysed and documented in this section.
During TR3 GMV operates through the OBC the four FCVs following a predetermined reference profiles. This profile was than executed by the Maxon engine that moves the FCVs and retrieved through its encoder to check if the command was really executed and with which error through the FCV.
After analysing the results two major problems were identified: 1) a bias existing among the commanded signal and the one retrieved through the encoder, 2) a delay among the sent command and the executed one.
Both the problems have been solved and the run 2 showed positive results

vi. TR3B (firing)
Scope:TR3b is final point of TR3 firing sequence, as it involves the step from water simulant (TR3a) to real propellant (87.5% H2O2).
The goals of the test were:
1. To measure motor performances, thrust and the moments transferred to structure (the lander was attached to the SPARTAN test fixture using adapter and appropriate 6-DOF load cell)
2. To test the control avionics (OBC, Epos controllers), validating the actuation capability, the navigations sensors (Inertial platform, LIDAR) and the monitoring sensors (pressure sensors) in flight like configuration.
3. To perform an end-to-end firing test, verifying:
• Full chain delay: command to actuators
• Synchronicity in terms of ignition and throttling

Conclusion: TR3b was successful for the following reasons:
• The system compliancy for operation with high grade H2O2 was fully demonstrated.
• OBC was able to operate the four engines in a fully synchronized way, applying the selected thrust profile, the same for all engines.
• All equipment worked as expected and all procedures for filling, firing and draining were proven to be applicable, safe and reliable.
• All four motors have been fired: pressure, thrusts and torques generated by the complete system were measured. Thrusts are well in line with expected theoretical values, despite accurate measurement of single motor thrust was not possible (due to derivation of these values from combined torques at the load cell).
• Mass flow rate and mass H2O2 consumption seems to be well in line with extrapolations from theoretical profile.
• The test was opportunity to test LIDAR and IG-500N integrated in a flight like configuration.
The heaters were not able to reach the desired temperature of 200°C, stabilizing at a maximum of about 75°C.

vii. TR4A  Open Loop Dynamic System Firing Test (quasi-unrestricted)
Scope: The TR4 is carried out quasi non-restricted movement drop tests to fully verify the Landers capability in maintaining its altitude and attitude during a lift off, hovering and controlled landing, before moving to the free flight.
Even if different with respect to the final flight mission statement, the TR4 is defined in a way the all the lander functionality are tested in an end-to-end closed loop scenario.
The TR4 is approached in two steps:
• TR4-A open loop
• TR4-B closed loop

The profile is the same, but, for safety reasons, it has been decided to carry out the first run with the navigation functions in the operator hands. This approach was selected in order to be able to validate the soft landing System capabilities minimising all the risks to damage the lander itself.
This is, in fact, needed for the flight test and any further eventual development testing.
Starting point that made it possible to trigger the switch to TR4 test phase was the successful conclusion of TR3 sequence, leading to the following results and lessons learnt:
• System compliancy for operation with high grade H2O2 demonstrated
• Procedures for filling, firing and draining proven to be applicable, safe and reliable.
• Mass flow rate, H2O2 consumption and thrust in line with theoretical values
• Time to ignition to be ensured by means of igniters, to be implemented for TR4
• LIDAR and IG-500N tested in flight like configuration.
o Mechanical decoupling needed for inertial platform. SAD to be implemented for TR4
o LIDAR blindness due to monopropellant phase to be solved by introduction of igniters

TR4 was the final ground test sequence for SPARTAN system validation and focused in particular on validation of avionic hardware and software. The goals of the test were:
• Validate the avionic hardware and check it functionality in a flight representative test environment.
• Validate GNC algorithms for attitude control.
• Test the igniters and their effect in terms of reduction/elimination of monopropellant phase (time to ignition).

For TR4 the SPARTAN system was brought to a flight like configuration, including all devices needed for an hovering sequence:

For TR4 the complete avionics configuration has been employed. The avionics system, provided as a turn-key GNC set, is made of the following main functional system blocks:
• PDU box including: Four Ni-MH batteries, actuators Switches DC/DC converters for each output line, EPOS controllers for FCV, DAU for voltage and current analog input (sensors).
• OBC box including: OBC MB, MMU, WiFi/LAN TM module, Relay board and PC/104 CAN-Interface board.
• Central pillar cartridge bracket equipment including Inertial Navigation System (INS) and Laser altimeter (LIDAR).

Each block is enclosed in a stand-alone, dedicated case with all mechanical/electrical interfaces needed. Furthermore, the following stand-alone components are mounted:
• Cameras
• Antennas for both GPS and Wi-Fi telemetry

Specific test sequences, implemented in a coded process, were prepared for bot TR4A (Open Loop) and TR4B (Closed Loop) Profiles:
• System health checks and definition of GO/NO-GO criteria have been defined as flow charts, implemented as code and tested.
• Dedicated telemetry synoptic have been developed and tested to keep under control
• All equipment is controlled by on-board software by means of the on board avionics. Sensors output is read and stored in MMU.

The lander included legs and had to be constrained in order to avoid safety issues during hovering. For the purpose, the vertical test fixture was modified including guidance cables allowing total freedom on vertical axis only (quasi-unrestricted)

Achievements: A malfunction to FCVs actuator led to no ignition of two motors and didn’t allow to perform the hovering sequence. However the test was successful in providing the following results:
• Engine igniters successfully tested both separately and at system level.
• Legs and damper installed. Ballast mass, to recover COG position, installed.
• Avionics: actuators, sensors and telemetry are now in the loop, tested and proved. The telemetry provides correct information and records data on the OBC.
• Code and algorithms validated with E2E approach.
• Avionics Procedures validated.

viii. TR4B  Closed Loop Dynamic System Firing Test (quasi-unrestricted)
It is planned to run this test once the FCV failure investigation will be closed and the valve fully recovered. This is not a matter of concern, because the valves themselves demonstrated successful behaviour several times at different operating level:
• Valve level
• Engine level
• System level

ix. Demonstration Flight Test
It is planned to be carried out once TR4b can give green light and the GNC is verified accordingly to the TR4b test results. The test will be carried out at a CZ airfield using a heavy weight helicopter
6. Objectives Achievements

Objective 1: The engine design, specific for throttling functionality
• The Engine design is defined and justified within the deliverable D 4.1
• The engine throttling performances have been demonstrated through TR1, TR2A, TR2B.
• The relevant results are reported in the annexed documentation

Objective 2: The oxidizer throttable device development
The Throttling Device was designed, analysed and tested in a laboratory environment with both Water ad Hydrogen peroxide. It was then installed on the engine and tested with hydrogen peroxide. Finally it was installed on the lander and tested with water for its final validation. The test results are reported in the deliverable D 2.5 ad D4.5 and in the annexed documentation

Objective 3: Develop the Landing test capabilities targeted to:
• Test bench
o achieved
• Test procedure
o achieved

Develop the Landing test capabilities targeted to:
• Perform the soft landing, by damping the vertical acceleration
o TR3-B demonstrated end-to-end actuation capability at system level: synchronicity
• Maintain the Lander stability: will be finally demonstrated by the TR4-B (closed loop dynamic system firing test): sensors in the loop

Helicopter Test Description
The test specification has been already established and reviewed.
The test is defined to carry out the specified mission statement:
• Release altitude q0: ≥ 170 m
• Free fall distance: ~ 50 m (target speed at ignition is ~ 30 m/s)
• Engine ignition altitude q1: ≥ 120 m

Failure/Anomaly Recovery case:
The mission recovery is automatically driven by the FDIR, implemented by the OBC.
The Recovery system is triggered by the following events, which will be eventually detected starting from q1. According to the FDIR, the parachute should be ejected and the landing will occur at 6.5 m/s maximal impact velocity.
Safety aspects have been treated and validated by NAMMO, the Fire Brigade and by the hydrogen peroxide provider.

The selected helicopter is Mil Mi-8 Medium Utility Helicopter. During the SPARTAN project there were tested in real conditions two types – Mi-8 and light utility helicopter Bell 407. Based on evaluation of SPARTAN dummy structure lifting and hovering tests, there was selected heavier and more stable Mi-8.

The Mil Mi-8 is a twin-engine medium utility helicopter developed by the Soviet manufacturer Mil OKB, today MIL Moscow helicopter plant, JSC (Russia).

For the final test there was selected Heliteam CZ Ltd. Company. Main reasons were – only operator with availability of Mi-8 heavy helicopter equipped with enough long electrically actuated hook for safe operation, the company very wide experience with heavy cargo transportation with precise positioning for delivery, experience with scientific experiments with academy of science.

Heliteam CZ Ltd.
Heavy lifting company
Toužimská 588/70, Kbely Praha 197 00 Czech Republic

The selected test-field is former military jet-fighters airfield close to Milovice city, Czech Republic. Airfield parameters:
Code: LKML - Milovice radio 125,825
Operations permitted: VFR - day
Location: 2,5 km NE Milovice, CZ
Coordinates: N50°14’10” E014°55’25”
Runway 09/27 096/276 2500x80m concrete
ELEV [m]: 199, ELEV [ft]: 653

Test Preparation Sequence

Test seq. Test description Responsibility
1 Inspection VUT
2 Lower Panels Installation VUT
3 Legs and dampers Installation & test VUT
4 Pressure test VUT
5 BRS Installation VUT
6 Upper Covering panels Installation VUT
7 Mass measurement VUT
8 Transportation to Air Field (including packaging) VUT

Test Preparation Sequence at Air Field

Test seq. Test description Responsibility
1 Inspection VUT
2 Health check configuration VUT
3 Engines Installation (TBC) NAMMO
4 H2O2 transfer tank status monitor NAMMO
5 H2O2 loading apparatus health check NAMMO
6 Battery Charging GMV
7 Avionics health check GMV
8A GNC sensors health check GMV
8B Magnetometer Calibration activity GMV
9 Telemetry check GMV
10 Oxidizer system status check VUT
11 H2O2 loading NAMMO
12 Gas Pressurization VUT
13 Final Functional check TBD
14 Radio Link Check VUT

Overall General Achievements
The capability of managing complex project with relatively low budget and reduced time frame: a very high level cooperation attitude has been demonstrated by all the partners:

• The capability of designing high level performance throttling devices;

• The capability of designing throatable hybrid engine;

• The capability of simulating in a lab the hydrogen peroxide with water and oxygen only;

• The capability of finely coding the behaviour of a hybrid engine combustion chamber;

• A reusable and improvable soft landing test vehicle;

• Ground test facilities and procedure;

• The capability to manage hydrogen peroxide at both ground and vehicle level;

• A throttling hybrid engine system that could have several additional applications than the designed lander one.

7. Conclusions & Lessons Learned
The SPARTAN SYSTEM has been fully developed, integrated and tested so far the equipment and S/S is concerned.

The system functionality has been verified up to TR4, enabling the validation of:
o The avionic hardware and check it functionality in a flight representative test environment.
o GNC algorithms for attitude control.
o Test the igniters and their effect in terms of reduction/elimination of monopropellant phase (time to ignition).

The system compliancy for operation with high grade H2O2 was fully demonstrated.

All four motors have been fired simultaneously: pressure, thrusts and torques generated by the complete system were measured.

Lessons Learned

• Landing gear design and performance correlation.

• Proper decoupling of INS IG-500N from structure mechanical environment: use of SAD has been introduced.

• Communication protocol improved to avoid reading errors found during nominal engine operation.

• Defined proper procedure to dry the pipework before hydrogen peroxide loading, to avoid prolonged start up sequence.

Project Statistics

• Number of delivered documents : 45 (~1/month)
• Number of Issued documents: 120 (~3/month)
• Number of Milestones: 9 (~1/5month)
• Number of meetings (including teleconf): 39 (~1/month)
• Number of tests: 76 (~2/month)
• Number of videos: 26 (~1/2month)
• Number of contacts on web: 3042 (~72/month)
• Number of papers, articles and presentations: 31 (~1/1.5month
• Number of Patents: 1

Potential Impact:
Description of the Potential Impact
The SPARTAN project developed and realised a Landing Demonstrator that could be used, together with the test validation facilities, for low cost, ready to use for technology enabling and ground validation:
The below list highlight those technologies ESA intends to advance through and that they can be supported by the SPARTAN achieved capabilities:
• Precision Landing and HDA GNC and avionics;
• Advanced GNC technology for ascent;
• Chemical propulsion systems for planetary landing and ascent;
• High Thrust Engine for improving Mars insertion;
• Mars Precision Landing technologies.
• Vision based landing system
This project was born by the need in Europe of throttling propulsion able to support a soft and precision landing on Moon and mars primarily, but not limited to the above.

Mission Scenarios
The Space Exploration policy and the new Mission Scenarios, in fact, are driven by the following principles:
o Capability Driven Framework: Follow a phased/stepwise approach to multiple destinations.
o Exploration Value: Generate public benefits and meet exploration objectives.
o International Partnerships: Provide early and sustained opportunities for diverse partners.
o Robustness: Provide for resilience to technical and programmatic challenges.
o Affordability: Take into account budget constraints.
o Human-Robotic Partnership: Maximize synergy between human and robotic missions.
Current affordability constraints and the available budget allow to focus investments up to 2020 on implementing the running programmes (ISS and ExoMars/MREP), while preparing focused engagement in the international space exploration endeavour for the post 2020 era, strongly leveraging on international cooperation opportunities.
ESA MREP (Mars Robotic Exploration Preparatory Programme) approach is to consider a Mars Sample Return (MSR) mission as a long-term objective and to progress step-by-step towards this objective through short and medium-term MSR-related technology developments, which are validated during intermediate missions and developing Long Term enabling technologies, such as Novel Power Systems (NPS) and Propulsion engines.
Within this frame there are two proposed candidate intermediate missions:
• The Mars Moon Sample Return mission, also called PHOOTPRINT, aiming at returning a sample from a Mars moon, nominally Phobos.
• The Network Science mission, also called INSPIRE, aiming at delivering a network of probes on the Mars surface to perform simultaneous seismic and meteorological measurements.

Mission Road Maps
The primary focus of the International Architecture Working Group (IAWG) during 2011 was to develop and refine the ISECG mission scenarios.

Six mission scenarios have been initially investigated. Three of them were eliminated due to high technological challenges, constrained budgets, high risk, and insufficient mission opportunities.
The three remaining scenarios were evaluated by the IAWG. After further consideration, the number of scenarios was reduced from three to the following two: “Asteroid Next” and “Moon Next”.
The IAWG then focused its efforts on concepts for displaying the key features of the two remaining scenarios to effectively communicate each strategy at a summary level.
They differ primarily with regard to the sequence of sending humans to the Moon and asteroids; each reflects a stepwise development and demonstration of the capabilities ultimately required for human exploration of Mars.
A paper titled “ISECG Mission Scenarios and Their Role in Informing Next Steps for Human Exploration Beyond Low Earth Orbit” was presented at the 62nd IAC in 2011.
According to the changed global scenario, the Roadmap has been revised and the ESA Roadmap, accordingly, is derived from and fully consistent with the ISECG Global Exploration Roadmap.
The above shows that ESA would participate in two strategic areas:
• The first is related to gaining access to the lunar surface: on the capability to transport crew to the vicinity of the Moon with NASA and on the capability to access and return from the lunar surface with Russia. This opens perspectives for attractive and sustained roles in lunar exploration.
• The second is related to the participation in the first mission to return sample from Mars.
To support these two objectives ESA is planning to put on the critical path the following criticalities:
• The development of a staging post in cis-lunar space. Such staging post advances deep space exploration and lunar surface exploration capabilities concurrently.
• A human-assisted sample return, re-usable lunar lander.
• The development of an international lunar landing capability.
The gradual evolution of capabilities will allow in a step-wise approach to implement more and more demanding and complex exploration missions, with complexity being primarily a function of the energy requirement, overall mission duration and the constraints/ limits for emergency returns of Astronauts.
SPARTAN, as lander demonstrator, will play a significant role supporting this kind of implementation and development approach, together with the exploitation of synergies between human and robotic missions and the definition of an integrated mission concept.

Small Launcher Application
The idea of a dedicate launcher for very small payloads is discussed in US and Europe by many institution (ESA; DARPA, CNES among others). In fact, the recent multiplication of micro satellite platforms and the increasing success to the associated applications allow us to think about the interest of a dedicated launch system for small payloads:
• Numerous successful technology experiences based on microsats, for Science or Defense with gradual increase in instrument performance.
• Constant improvement of quality / price ratio.
• Rebirth of interest for constellations (Rapid eyes, Orbcomm2, numerous projects worldwide…) and formation flying.
• Increasing number of operational applications accessible: communications, intelligence gathering, early warning, space surveillance, different type of observation, etc.
• Evidence of the vulnerability of the big space systems.
• Increasing interest for the « Responsive Space » approach in the USA and other countries (China, etc.), which prefer small size in order to reduce global costs et delays, and facilitate the implementation of new technologies.
The constellation of very small satellites (1-40 kg) is dominated by CubeSats. A CubeSat is a type of miniaturized satellite for space research that usually has a volume of exactly one litre (10 cm cube), has a mass of no more than 1.33 kilograms, and typically uses commercial off-the-shelf electronics components. Beginning in 1999, California Polytechnic State University (Cal Poly) and Stanford University developed the CubeSat specifications to help universities worldwide to perform space science and exploration. CubeSat initiatives are growing each year and advanced satellites made up by adding several CubeSats have already been proposed. As an example a 6 units CubeSat could be capable of earth imaging with medium resolutions (5-7 meters per pixel).
The nano/microsatellite market has grown considerably with the adoption of the CubeSat standard, microelectronics and other technology development, entrance of new developers, new government programs, and furthering of applications. Projections based on announced plans of developers indicate 121-188 nano/microsatellites requiring launch in the year 2020.
Nano/Microsatellite CAGR (Compound Annual Growth Rate):
• Historical average growth of 8.6% per year over the last 12 years (2000-2012)
• Announced Dataset average growth of 16.8% per year over the next 7 years (2013-2020)
• Optimistic Dataset average growth of 23.4% per year over the next 7 years (2012-2020)
Historical and announced future data set suggests that the average number of nano/microsatellites launched per year triples with every five year period (2001-2005, 2006-2010, and 2011-2015)
The current market for very small satellites (<500 kg) is more active in the microsats range (40-150 kg). Although over the last few years there has been a strong increase in demand for launching of these satellites, for which more and more platforms are available. This increase in number might also be accompanied by an increase in satellite masses (increase from 150 to 200 kg, or even more) in order to optimize mission yields.
Nano/microsatellite (1-50 kg) development continues to be led by the civil sector, but the defence/intelligence community is showing increased interest and involvement. Applications for nano/microsatellites are diversifying, with increased use in the future for science, Earth observation, and reconnaissance missions.
The smallest launchers available today have payload capabilities in the hundreds of kilograms range for LEO and launch costs of 10-20 millions. Nowadays the best option (as seen from the number of microsat launches) is using the Indian PSLV and Russian small launchers even if the Indian PSLV is still too big and expensive for a responsive, affordable micro/nano satellite market.
Russia’s current inventory of small launch vehicles provides a range of services for the launches of various classes of small satellites at relatively low launch prices. All the inventory’s launchers, excluding “Cosmos-3M”, are originating from ballistic missiles for which the guaranteed lifetimes are near to expiration. Russia should be able to substitute its current small launchers with the Angara 1 launch vehicle. However this launch vehicle is relatively large and cannot effectively compete with the actual “doped” low price of ex-soviet ballistic missiles.
Aside from the new growth of a small satellite market, the launch options are not capable of addressing to it. The two main reasons of the lack of competition of classical micro launch vehicle are:
• Scale effect: the bigger the launcher the less expensive is the cost/kg of the payload. The scaling ratio suggests that is much easier to design a big launcher with very small inert mass fraction than a small launcher with the same low inert mass fraction. Interfaces, valves, piping and other items do not scale with the propellant and this gives the well know result that the more the propellant that is needed, the less the inert mass that is required to operate the launch system. The launch cost trends are shown in table below.
• Fixed cost such as expensive ground facilities required to assemble the launcher and the payload, to operate the launcher, to guarantee a safety zone and to manage the propellants.
One solution of the problem is to reduce at minimum the cost connected with the propulsion system.
The first step the use of an Airborne Launcher: it consists in the use of a re-usable, existing and very reliable first stage such as a military aircraft of cargo aircraft that shall deploy the small satellite launcher at an altitude of roughly 10 kilometres.
As suggested also by DARPA in a study proposal issued in year 2012 for the similar application (responsive small satellite launcher vehicle with airborne first stage), hybrid propulsion technology is a candidate to overcome the technical issues still holding. Unlike the mature propulsion technologies (such as the solid and liquid rocket systems), which could only result in small incremental improvements, hybrids do have the potential to become a game changing propulsion technology which could potentially lead to significant cost savings in a relatively short period of time, enabling safe handling of the launcher as well.

Exploitation Strategy
In accordance to the SPARTAN exploitation plan, The main aspects of the exploitation strategy are:
• Support to system architecture studies/trade off/feasibility studies for space exploration mission, concepts and applications;
• Identification of other potential applications for the proposed hybrid throatable technology (i.e. ascent from planets, surface mobility, LEO small and low cost platform);
• Definition of a technology development, deployment and exploitation plan and of transfer processes;
• Definition/identification of project products implying final system sub-features, whole system (concept, know-how, expertise and relevant services) and of the royalty rights to the aforementioned products;
• Develop the basis for the business beyond the project for the products that are much closer to industrialization (definition of a business plan and of appropriate market strategies);
• Establish of a research and industrial network, enabling the exploitation of the hybrid engine technology with throttling capabilities and its further and continuous development, addressed to any possible kind of application.

Exploitation Guidelines
The SPARTAN development was driving for many technology acquisitions for each individual partner.
Here below a list of exploitation possibility, associated to the single partner is provided:

Lander demo vehicle, for landing performances ground verification purposes. There is a high interest of ESA in exploiting the already developed vehicle, for further ground landing verification tests, pending the system refurbishment.
• Future mission landing design and development
• Support to IXV re-entry FDS development and verification
• Support to EXOMARS 2018 Landing FDS
• Green propellant system management
• Green propellant tank

The developed SW has been intensively used at UNIPD for engine transient analysis, to verify unstable phenomena within the frame of MIUR research program (Italian Ministry of research and University)
Similarly for the gas generator and small scale test set-up, used, and planned to use it, to support research program on engine development of both H2020 proposal and MIUR programs

The space activity is currently under expansion with development of a new generation hybrid rocket motors for ESA and the Norwegian Space Agency, as well as development of green propulsion systems with small thrusters utilizing H2O2.
Development of LEO small and low cost platform, in order to ease the access to space.
Possibility to reuse the developed test bench for ground demonstration purposes: static and dynamic (quasi-restricted)
Use the H2O2 filling equipment for monoprop systems as well.

MOOG Bradford
Any kind of liquid throttling valve request, coming from either propulsion system functions or any kind of liquid delivery systems.
At the present time, the achieved capabilities, are going to be exploited within the frame of further H2020 proposal and to further implement the SPARTAN demonstrator, according to the ESA interests, led by TAS-I, acting as the SPARTAN system prime.
The SPARTAN flow control valve has been added as a new product to the portfolio of Moog Bradford. It is also be the first product specifically intended for use with hydrogen peroxide, where existing products have been focused towards gaseous or other liquid propellant applications. At the time being Moog Bradford started actively marketing the new hydrogen peroxide flow control valve, since the development in frame of SPARTAN has been completed.

1st ITT was focused on Marco Polo subscale model dynamic stability free fall tests via helicopter drop test (prepared with TAS-I, Martins Sudars).
2nd ITT was focused on similar topic, re-entry vehicle subscale model stability characterization via stratospheric baloon drop tests (submitted with GMV).

A new concept of for a Regression Rate Sensor (fuel diagnostics) has been invented.
The SPLab-POLIMI has worked at developing this kind of measurement and at porting it in a configuration that is suitable for in-situ installation inside a fuel grain, before the casting. The electronic circuit is reproduced in a circuit board with a specific design of electronic components and of sensor head.

The Sensor is currently under patent process evaluation. This sensor has been included in an on-going H2020 proposal – hybrid engine.
In the near term, it is planned to use the sensor to evaluate the regression rate o a solid rocket motor, in cooperation with a major European Company.

On what concerns with Spartan main results presentation, GMV prepared presentations and Videos of the main experiments under its responsibility that have been distributed and presented among the GNC community in Europe. During the last ESA Guidance Navigation and Control Conference (held in Porto last June) GMV has been very active in presenting Spartan main results to different entities working worldwide in this area through Spartan dedicated videos showed up at the GMV stand. Presentations and videos was shown to attendees coming from ESA TEC-ECN section, to the head of Control Division, to the head of GNC Section and to personnel coming from DLR and Draper Laboratory. Feedback was very positive from all the community that shows interest in the project and in a possible continuation.
GMV along with TAS-I intends to insist with the aerospace community in a possible exploitation of this technology and the available flying platform. Contact with main player in ESA has been already established and possible exploitation of the platform and throatable technology already defined.

Major Achievements:
• Impact analysis of a complex system
• Legs design
• Damper and damping requirements definition.

Current and forecast exploitation:
• Provide support to TAS (either the Italian or the French branch), to develop lander structure, or part of it, applying the developed impact analysis tools.

List of Websites:
Project Website:
Project Coordinator: Thales Alenia Space Italia S.p.A.
• Enrico Gaia – Coordinator (
• Mario Pessana – Technical Coordinator (

Partners – technical contacts:
• Università di Padova – Daniele Pavarin (
• Nammo Raufoss AS - Jan Erik Rønningen (
• Moog Bradford - Harrie Leenders (
• Brno Technical University (VUT) – Robert Popela (
• Politecnico di Milano - Filippo Maggi (
• GMV - Emanuele Di Sotto (
• INGEMECA Groupe STUDIEL - Valentine Stasse (

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