STOCKHOLM SWEDEN 2019,
Design, assembly and
commissioning of a flexible test bench for propulsion system components
GABRIELE GUERRA
KTH ROYAL INSTITUTE OF TECHNOLOGY SCHOOL OF ENGINEERING SCIENCES
Sedan f¨oretagets begynnelse 2011 har OHB Sweden bem¨otts av ett st¨andigt ¨okande behov att utveckla b˚ade kemiskt och fullst¨andigt elektriska satelliteframdrivnings- system. En av de st¨orsta kostnaderna, b˚ade i tid och i pengar, f¨or s˚adana processer
¨ar projektens verifikationssteg. Detta verifikationssteg best˚ar av en kombination av processor som tillsammans visar att produktens krav specifikationer uppfylls p˚a ett tillfredst¨allande under hela produktens livscykel, vilket g¨ors fr˚an minsta komponent till den f¨ardigbyggda rymdfarkosten. D˚a antalet projekt hos f¨oretaget
¨okat explosionsartat beh¨ovdes en innovativ metod f¨or att minska arbetsbelastning under projektens verfikationssteg.
Detta examensarbete beskriver utvecklingen av en ny typ av testb¨ank f¨or kvali- fikation och godk¨annande av ett framdrivningssystems olika delar, som till˚ater stor testm¨ojlighet med minimal testb¨anksmodifiering. En nyutvecklad testb¨ank skulle inneb¨ara stora framdrifter hos OHB Sweden d˚a direkta och indirekta kostnader kan kapas av f¨orenklade testuppst¨allningar, d˚a tidigare testb¨ankar kr¨avde komplex is¨ar- och ihopplockning av speciella testb¨anksvarianter. Testb¨ankens huvudsakliga m˚al ¨ar att den ¨ar flexibel f¨or varierade testuppst¨allningar. M˚alet uppn˚as genom att de relevanta komponenterna monteras i s¨arskilda konsoler som i sin tur kan glida fritt l¨angs med aluminiumr¨als. Detta s˚a att anv¨andaren slipper montera av eller p˚a de specifika konsolerna vid olika testuppst¨allningar.
Since its foundation in 2011 and among other things, OHB Sweden has been appointed to fully develop, from the design to the assembly, a yearly increasing number of propulsion systems, ranging from chemical propelled to fully electrical spacecraft. One of the major contributor to the cost of this process, both in terms of money and time spent, is the verification phase, a combination of processes that sets out to demonstrate that all the applicable requirements are met and to prove that the system is capable of fulfilling its objective during the mission lifespan.
The increasing number of projects running simultaneously led to the need for an innovative, smart solution to reduce the overall efforts spent on verification.
The thesis work described in this document addresses the development of a new test bench that allows carrying out qualification and acceptance tests on differ- ent components or sub-assemblies of propulsion systems while requiring minimum setup modifications. The achievement of this goal would result in great benefits to OHB Sweden, decreasing both the direct and indirect costs of the testing activities.
Indeed, the current baseline approach consists of developing and assembling a new setup every time a test is scheduled. Flexibility of the bench is a major goal: it is achieved by mounting relevant components in custom design brackets that can in turn slide on aluminium rails. This ensures the possibility of adding, removing or replacing components without the need of disassembling the entire setup.
1 Introduction 15
1.1 Problem definition . . . 15
1.2 Verification process . . . 16
2 Spacecraft Subsystems 21 2.1 General description . . . 21
2.2 Propulsion system . . . 23
2.2.1 Cold gas system . . . 25
2.2.2 Liquid system . . . 25
2.2.3 Electric propulsion system . . . 28
2.3 Typical components . . . 28
2.3.1 Valves . . . 29
2.3.2 Pressure Transducer . . . 32
3 Test Bench Requirements Definition 37 3.1 OHB references . . . 37
3.1.1 Euclid . . . 37
3.1.2 Electra . . . 40
3.2 Test matrix . . . 41
3.3 Test description & requirements . . . 42
3.3.1 Leak tests . . . 42
3.3.2 Flow rate & Pressure drop . . . 49
3.3.3 Water hammer . . . 52
3.3.4 Slam start . . . 54
3.3.5 Noise on regulated pressure . . . 55
3.3.6 Proof pressure . . . 57
3.3.7 Pressure cycling . . . 57
3.3.8 Bubble point . . . 58
3.4 Test bench requirements specification . . . 60
4 Test Bench Design and Integration 65
4.1 Approach . . . 66
4.2 Schematic . . . 66
4.3 Components requirements and selection . . . 68
4.4 Design for flexibility . . . 74
4.5 Ground support equipment (GSE) . . . 75
4.5.1 Helium leak detector . . . 75
4.5.2 Pressure panel & Pressurization system . . . 76
4.5.3 Vacuum chamber . . . 77
4.5.4 Vacuum pump . . . 78
4.6 Acquisition system . . . 78
4.7 Assembly . . . 81
4.8 Schematics . . . 83
5 Commissioning Tests and Conclusions 91 5.1 Test Setup . . . 91
5.2 Leakage checks . . . 93
5.2.1 Sniff test . . . 94
5.2.2 Pressure decay test . . . 94
5.3 MIV flow rate vs pressure drop test . . . 97
5.3.1 Test Setup and Procedure . . . 97
5.3.2 Measurements and success criteria . . . 97
5.3.3 Test Results . . . 97
5.4 Conclusions and future works . . . 99
Appendices 105 Appendix A - 3D models 107 A.1 Valves brackets . . . 107
A.2 Inlet cross fitting support . . . 109
A.3 Pneumatic actuator support . . . 110
A.4 Differential pressure transducer support . . . 112
A.5 Solenoid valve support . . . 113
A.6 Mass flow meter support . . . 114
A.7 Rail and sliding nut . . . 115
A.8 Valve mounting example & Vertical alignment . . . 116
Appendix B - Conversion formula 119
2WV 2 Way Valve 3WV 3 Way Valve AI Analog Input
AIV Assembly, Integration and Verification AO Analog output
AOCS Attitude and Orbit Control System ASD Amplitude Spectral Density
BFSL Best Fit Straight Line BOL Beginning Of Life
C&DH Command and Data Handling System CDR Critical Design Review
Cp specific heat at constant pressure Cv specific heat at constant volume CV Check Valve
DAQ Data Acquisition System DI Digital Input
DO Digital Output
ECSS European Cooperation for Space Standardization EGSE Electrical Ground Support Equipment
EOL End Of Life
EP Electric Propulsion
EPPS Electric Propulsion Subsystem EPS Electrical Power System
FAR Flight Acceptance Review FCV Flow Control Valve FDV Fill and Drain Valve FS Full Scale
FVV Fill and Vent Valve
g0 standard acceleration of gravity γ heat capacity ratio
GSE Ground Support Equipment HET Hall Effect Thruster
HP High Pressure HV Hand Valve IPA Isopropyl Alcohol
Isp Specific Impulse IV Isolation Valve LP Low Pressure LV Latch Valve
˙
m mass flow rate mm molar mass
MDP Maximum Design Pressure
MEOP Maximum Expected Operating Pressure MFM Mass Flow Meter
MIV Manual Isolation Valve MLI Multi Layer Insulation
MPFA Micro Propulsion Feed Assembly MPR Mechanical Pressure Regulator MPS Micro Propulsion System
MTA Micro Thruster Assembly NC Normally Closed
NO Normally Open NTO Nitrogen Tetroxide NV Needle Valve
NVbp Needle Valve of bubble point line OR Orbit Raising
PDR Preliminary Design Review PPU Power Processing Unit PR Pressure Regulator
PT(s) Pressure Transducer(s) PTR Post Test Review
PV Pyro Valve Ql leak rate
QR Qualification Review ROD Review of Design
RTD Resistance Temperature Detector scc standard cubic centimeter
SKM Station Keeping Manoeuvres SV Solenoid Valve
T Thrust
TC Thermocouple
TCM Transfer Correction Manoeuvres
TCS Thermal Control System TPRO Test Procedure
TRB Test Review Board
TRL Technology Readiness Level TRPT Test Report
TRR Test Readiness Review TSPE Test Specification
TTS Telemetry and telecommand System VLP Very Low Pressure
WT Wall Thickness
2.1 Spacecraft subsystems classification . . . 22
2.2 Summarization of system propulsion options . . . 24
2.3 General architecture of self-contained energy propulsion system . . 25
2.4 General architecture of externally supplied energy propulsion system 25 2.5 Typical schematic of a cold gas propulsion system . . . 25
2.6 General schematic of liquid systems. . . 27
2.7 Pressurization system options, from [5]. . . 28
2.8 Schematic of ion thruster . . . 29
2.9 Hand Valve example . . . 30
2.10 Needle Valve example . . . 31
2.11 Normally Closed solenoid valve . . . 31
2.12 Check Valve example . . . 32
2.13 Working principle of an absolute pressure transducer . . . 33
2.14 Working principle of a gauge pressure transducer . . . 34
2.15 Working principle of a differential pressure transducer . . . 35
2.16 Pressure regualtor schematic . . . 36
2.17 Filter example . . . 36
3.1 Euclid artistic representation . . . 38
3.2 Euclid MPS . . . 39
3.3 Electra artistic representation . . . 40
3.4 Electra booms configuration . . . 40
3.5 Electra booms kinematic . . . 41
3.6 Molecular flow path of a single molecule . . . 45
3.7 Laminar flow inside a pipe . . . 45
3.8 General schematic of the internal leak test. . . 47
3.9 General schematic of the external leak test. . . 48
3.10 General schematic of the mass flow and pressure drop test. . . 50
3.11 Pressure regulator test using gas medium . . . 51
3.12 Water hammer peaks in test with different initial tank pressures . . 53
3.13 Valve water hammer test . . . 55
3.14 Pressure transducer water hammer test . . . 55
3.15 Pressure regulator slam start test . . . 56
3.16 Pressure transducer slam start test . . . 57
3.17 Example of pressure noise ASD mask . . . 58
3.18 FFT transformation of a signal, from time to frequency domain . . 59
3.19 General schematic of the proof pressure test . . . 59
3.20 General schematic of the pressure cycling test . . . 60
3.21 General schematic of the filter bubble point test . . . 60
4.1 Bench design approach . . . 66
4.2 Bench sliding brackets concept . . . 75
4.3 Helium detector . . . 76
4.4 Pressure panel and pressurization system . . . 77
4.5 OHB vacuum chamber . . . 77
4.6 Vacuum pump . . . 78
4.7 NI cDAQ-9172 . . . 79
4.8 Architecture of the acquisiton system . . . 80
4.9 Example of user interface . . . 80
4.10 Test bench assembly . . . 81
4.11 Test bench assembly, top view . . . 82
4.12 Test bench schematic, version 0 . . . 83
4.13 Test bench schematic, version 1 . . . 84
4.14 Test bench schematic, version 2 . . . 85
4.15 Test bench schematic, version 3 . . . 86
4.16 Test bench schematic, version 4 . . . 87
4.17 Test bench schematic, version 5 . . . 88
4.18 Test bench schematic, version 6, only gas part . . . 89
5.1 Test setup for bench MIV test. Bench commissioning . . . 92
5.2 CAD model of the MIV mounted in its bracket . . . 93
5.3 Drawings of the MIV subjected to the acceptance fluidic tests . . . 93
5.4 Agilent G3388B leak detector . . . 94
5.5 Leak check in MIV fitting . . . 95
5.6 Pressure decay test result . . . 96
5.7 Flow rate vs. pressure drop test results . . . 98
5.8 Flow rate vs. pressure drop test results, zoom on the initial transient with differential pressure peak. . . 99
1.1 Verification stages and related milestone . . . 17
1.2 Verification levels . . . 17
2.1 Characteristics of propulsion options . . . 24
2.2 Valve types and functions . . . 29
3.1 Fluidic tests . . . 42
3.2 Test matrix . . . 43
3.3 Joule-Thomson Assessment, Electra MPR testing phase B1 . . . 51
3.4 Choked flow condition calculations, Electra MPR testing phase B1 . 53 3.5 Euclid filter pressure cycle specification . . . 58
3.6 Test bench requirements . . . 61
4.1 Schematic change log . . . 67
4.2 Bench components requirements . . . 69
4.3 Selected components . . . 72
4.4 Helium leak detector specifications . . . 76
5.1 Commissioning test, components specifications . . . 91
5.2 Pressure decay requirement vs. result . . . 96
5.3 Results of the MIV flow rate & pressure drop . . . 98
Introduction
This chapter first states the objective of this thesis work and the constraints to which it is subjected. Then it introduces the reader to the verification process, to which the testing activities belong to, briefly describing the different verification stages and levels in light of the ECSS1 standards [7],[8].
1.1 Problem definition
Maintenance activities on spacecrafts are usually not possible after they have been launched and placed on orbit thus even a single small defect or failure can lead to the loss of an entire mission. This translates into the need for intensive on-ground testing at all levels to verify that every part of the spacecraft fulfil its requirements and it is flight ready. Testing activities are essential but also very expensive and time consuming, heavily influencing the program cost and the overall schedule.
Moreover, tests are non-value-adding therefore an efficient solution has to be found to minimize their impact on the mission development. From this perspective, the goal of this work is to design a flexible solution for propulsion system components testing.
To date, when a test is scheduled one or more engineers are appointed to it and start elaborating the whole test structure. This implies a huge effort for designing the test set-up, procuring the needed supporting equipment and for preparing the
1European Cooperation for Space Standardization: it is an initiative established to develop a coherent, single set of user-friendly standards for use in all European space activities. The goal of ECSS standards is to improve quality of all elements of a space project. This goal is achieved by applying common standards for project management and for the development and testing of hardware and software. ECSS was created in 1995 [9].
data acquisition systems. Plus, the test set-up has to be assembled and commis- sioned before the test on the component can start. Weeks or even months are sometimes required for all the preparatory activities. Apart from the equipment cost, considering the hourly wage for an engineer it is clear how expensive the testing activity is. With this in mind, OHB Sweden has decided to put effort into finding a smart solution for decreasing the time spent for all the activities that pre- cede the test. Indeed, the goal of the thesis is to equip OHB propulsion laboratory with a fixed, flexible test bench on which quick characterization of components and sub-assemblies of propulsion systems can be carried out. The bench shall allow testing of different components, using different medium. The design of the bench is approached as the following optimization problem:
O1) maximize the type of test that can be carried out;
O2) maximize the number of components that can be tested;
O3) minimize the bench complexity;
O4) minimize effort in preparing the test setup when switching between different components / tests;
subjected to constraints:
C1) maximum bench dimensions 1.5 m x 1.2 m;
C2) budget limit;
C3) safety regulations of the laboratory.
1.2 Verification process
This section provides useful information about the verification process. Notions given here as acceptance and qualification stages will be of large use in the next chapters. Some words are then spent on the test specific reviews and on the related documentation.
Verification is the process whose objective is to demonstrate that the deliv- erable product meets the specified requirements and fulfils its intended purpose.
Verification is implemented all along the program at different stages and levels,
using different methods [7]. The verification stages are related to the project mile- stones and performed incrementally at different levels. The verification levels refer to the type of item under test, from a single component to the overall system.
Typical verification stages and levels are summarized in Table 1.1 and Table 1.2, both taken from [7].
Table 1.1: Verification stages and related milestone Verification Stage Milestone
Development PDR - CDR Qualification CDR – QR
Acceptance QR – FAR
Pre-launch FAR – Launch In-orbit Commissioning
Table 1.2: Verification levels Verification Level Example Component (or Part) Bearing
Equipment (or Unit) Valve
Subassembly Printed Circuit Board
Subsystem Thermal
Element Entire satellite
Qualification and acceptance tests will be carried out using the newly designed test bench therefore some details are given below.
• Qualification stage aims at demonstrating that the design is capable of meet- ing all the applicable requirements including proper margins. If the tested article passes all the qualification tests it is then referred to as “qualified”
for its intended use. Articles built to the same engineering data, following the same manufacturing processes and for the same intended use do not have to pass the qualification sequence anymore, they will only be subjected to acceptance tests.
• Acceptance stage aims at certifying that the article has been manufactured in agreement with the qualified design, is free from workmanship defects and acceptable for intended use.
Verification is executed by the following methods, normally complementary:
test, analysis (which include similarity), review of design and inspection. This list shows the order of precedence that, in general, provides more confidence in the results.
• Test consists in evaluating item performances and functions under represen- tative condition and environment.
• Analysis consists in performing theoretical or empirical evaluation using tech- niques as design analysis, modelling and computational simulation.
• Similarity consists in providing evidence that the item is similar to another one already qualified. The term similar intended as same design and function.
• Review of Design consists in using approved documents such as design spec- ifications and engineering drawings to show that the requirement is met.
• Inspection consists in visual determination of the item physical characteristics as conformance to document drawing or workmanship requirements.
From a test specific point of view, the entire test programme of a project shall be divided into blocks which depends mainly on the item under test. Typical test blocks are leak/proof pressure, thermal tests and functional & performance tests.
For each test block the following reviews are held, in the order here specified:
test readiness review (TRR), post test review (PTR), test review board (TRB).
The TRR is held before the start of each block to verify that all conditions allow proceeding with the test activity. For instance, test documents as test specification, procedure, schedule and pass/fail criteria are addressed at TRR stage. The PTR declares the test activity completed and allow the release of the item under test and test facility for further activity. During PTR the following are verified among others: - all test data were acquired, recorded and archived; - tests were performed according to the test plan, the test specification and the test procedures; - status of compliance of the item under test to the relevant requirement. The TRB is held to review all results and conclude on the test completeness and achievement of objectives. During a TRB all the test related documents are review and a lesson learned is written.
For each test block there are three major accompanying documents:
1. Test Specification (TSPE): it defines the test requirements, the purpose of the test, the test approach, the item under test and the set-up, the required GSE and instrumentation with their measurement accuracy and the pass/fail criteria, among others. The TSPE is used at each of the level specified in 1.2 as a basis for writing the relevant test procedure and test report. It shall be available at the test block TRR and on time to allow for the test procedure preparation.
2. Test Procedure (TPRO): it gives directions for conducting a test activity in terms of activity objective, description of the tested item and its configu- ration. It shall also include detailed step-by-step instructions for conducting the test activities accordingly to the test requirements specified in the TSPE.
The TPRO is prepared for each test to be conducted at each verification level specified in Table 1.2. During the execution of the test, the TPRO is filled in and it becomes the “as-run” procedure that is then included in the test report. Together with the TSPE, the TPRO shall be available at the test block TRR.
3. Test Report (TRPT): it describes test execution, results and conclusions in the light of the test requirements specified in the TSPE. For this purpose, among other things it contains the scope of the test, the test description, the as-run test procedures and the conclusions, with particular emphasis on the closeout judgment. It shall be available prior to the TRB.
Spacecraft Subsystems
This chapter describes the typical spacecraft architecture and division in subsys- tems. First, a brief overview of each subsystem is given, then focus is put on the propulsion one. Its functions are listed and the different type of propulsion options nowadays implemented on spacecraft are presented together with their general schematics. Eventually, the components included in the schematic are described from a mechanical and operational point of view.
2.1 General description
Every spacecraft may be divided into two principal elements, the payload and the bus. The former consists in the mission equipment or instrument, i.e. a telescope, the latter provides all the necessary resources and functions to the payload so that it is fully functional and it can achieve the mission objective. The bus in turn can be divided in subsystems, identified in Figure 2.1 and briefly described below.
• The attitude and orbit control system (AOCS) determines the spacecraft or- bital elements and orients it in desired directions despite the effect of dis- turbances that act during in-orbit operations. Sensors are used for attitude determination and actuators are used for control. In simple spacecraft, atti- tude control can be achieved with passive methods such as interactions with Earth’s magnetic or gravity fields. In complex spacecraft the AOCS mostly relies on reaction wheels and reaction control thrusters.
• The telemetry and telecommand system (TTS) provides communication be- tween the spacecraft and the ground operations center. Telemetry refers to
Figure 2.1: Spacecraft subsystems classification.
the signals sent by the spacecraft, telecommand to the signals that the ve- hicle receives from the ground station. Telemetry data consist of payload mission data and spacecraft housekeeping data, telecommand includes oper- ator commands to operate the payload and control the vehicle. Transmitters, receivers and one ore more antenna are used for the purpose.
• The command and data handling system (C&DH) distributes commands re- ceived from ground station to the other subsystems and controls the flow of information originated from sensors, instruments and payload formatting them for downlink or use by other subsystems. Data and commands are pro- cessed by an on-board-computer, coupled with the telemetry and telecom- mand system for information exchanging with ground.
• The electrical power system (EPS) provides, stores, distributes, and controls spacecraft electrical power. It consists of power source (i.e. solar panels, fuel cells), power storage (i.e. batteries), power conversion (i.e. voltage regula- tors) and power distribution equipments. It’s size is determined by the peak power consumption and the power requirement during eclipse. Moreover, degradation of batteries and solar cells at EOL shall be considered.
• The thermal control system (TCS) function is to maintain the temperature of the payload and of the components of the other subsystems within cer- tain temperature range during each mission phase. Minimum and maximum temperature shall be regulated as well as temperature gradients. Heat is generated both by the spacecraft components (thrusters, electronic devices, batteries) and by the environment (Solar radiation, Earth albedo). Dissipa- tion of the heat is only achieved through radiation, using both active (heaters,
refrigerator, others) and passive control (surface finishes, MLI, others). The balance between heat gained and radiated determines the spacecraft temper- atures.
• The structures and mechanisms system is designed to sustain the loads, to provide stiffness or stability and to mechanically supports all other subsys- tems. Also, it provides the interface with the launch vehicle. Mechanisms are included in all the others subsystems, providing essential task to their operation. Examples are the deployment mechanism for the solar panel and pointing mechanism for antenna. The major design goal is to achieve mini- mum structural mass considering all strength and stiffness requirements. It is important to note that even though the loads during in-orbit operation are small the spacecraft must be designed to withstand the high loads that develop during the launch and ascent phase. These include high acceleration and vibration levels.
• The propulsion system is detailed in the next section.
2.2 Propulsion system
The spacecraft propulsion system function is to provide the thrust necessary for:
- changing the orbital parameters (orbit transfer, interplanetary transfer);
- perform station keeping (orbit control);
- orient the spacecraft (attitude control).
State of the art propulsion systems for spacecrafts are summarized in Figure 2.2.
They can be classified in two macro groups [5],[10],[25]:
A) Propulsion systems with self-contained energy in propellants: the energy to produce the thrust is the chemical energy (or pressure) of the pro- pellant. Cold gas and liquid systems are included in this group. The general architecture, shown in Figure 2.3, consists of:
- storage system: the tanks that contain the propellant;
- valves, piping, etc.: connect the storage system to the thrusters and regu- late the propellant flow;
- electrical control unit: commands the electrically actuated valves and thrusters;
- thrusters: generate thrust.
B) Propulsion systems with externally supplied energy to propellant:
the energy to produce the thrust is supplied by an external source. Electric propulsion systems are included in this group. The general architecture is shown in Figure 2.4 and consists of:
- storage system: the tanks that contain the propellant;
- electrical power generator: i.e. fuel cells, batteries, solar panels;
- valves, piping, etc.: connect the storage system to the thrusters and regu- late the propellant flow;
- electrical control unit: commands the electrically actuated valves and thrusters;
- power processing unit (PPU): converts the electrical input from the power generator to appropriate voltage and current required by the thrusters;
- thrusters: generate thrust.
Table 2.1 provides comparison of the main characteristics of systems belonging to the two groups.
Table 2.1: Characteristics of propulsion options
Group Isp (s) T (N) Complexity
Cold Gas (s) A 50 0.05 Low
Mono-propellant A 250 0.5-20 Medium
Bi-propellant A 320 5-45000 High
Electric B 3000 0.001-0.2 High
Figure 2.2: Summarization of propulsion system options. Solid propellant are generally not used for spacecraft hence not included in the figure. It is worth noting that in the past years studies have been conducted on solid propellant micro thrusters, see [24].
Figure 2.3: General architecture of self-contained energy propulsion system.
Figure 2.4: General architecture of externally supplied energy propulsion system.
2.2.1 Cold gas system
Figure 2.5: Typical schematic of a cold gas propulsion system.
Cold gas system consists of a pressurized gas stored in the tank at high pressure (200-300 bar) and exhausted through a nozzle, with- out being subjected to combustion processes.
It is the simplest and cheapest options but with low performances, typically specific im- pulse Isp1=50 s and thrust T =50 mN. It is com- monly used for orbit maintenance and attitude control tasks. A typical cold gas system is de- signed as in Figure 2.5, Section 2.3 provides de- scription of the components.
2.2.2 Liquid system
In this system the propellant is stored in liquid form inside tanks and fed to the combustion
1Isp: parameter used for determining thruster performances. It is defined as the thrust force delivered per unit weight of propellant flow rate Isp= T
g0· ˙m [10].
chamber where it undergoes combustion processes. Thrust is then generated by expansion of the combustion products into a convergent-divergent nozzle. Here the thermal energy of the high temperature combustion gases is transformed into kinetic energy when the gases are exhausted to ambient at high velocity. Liquid systems are divided in mono-propellant and bi-propellant.
Mono-propellant systems are commonly used for orbit maintenance and atti- tude control but they are sometimes used also for midrange impulse requirements.
Typical performance values are Isp=250 s, T =0.5-20 N. Hydrazine (N2H4), that is the most common propellant, is decomposed in the thruster by a catalyst bed and the hot gases that generate are exhausted through a nozzle. It is a simple and reliable system but catalyst bed degradation usually limit its use for mission of long duration [17].
Bi-propellant systems are mostly used for orbit transfer and as primary propul- sion system on commercial spacecraft with high impulse requirements (i.e. geo- stationary communication satellites)[17]. They are characterized by high perfor- mances, typical values are Isp=320 s and thrust up to T =45000 N, but at the cost of increased system complexity, cost and weight. In bi-propellant system, a liquid fuel and an oxidizer (usually mono-methyl hydrazine CH3(NH)NH2 and nitrogen tetroxide N2O4) are separately injected in the combustion chamber of the thruster where combustion starts spontaneously in the moment the two liquid are mixed (hypergolic propellant). The combustion gases are then expelled at high velocity through a nozzle generating thrust. The general schematic for mono and bi-propellant system is shown in Figure 2.6. As mentioned previously, components explanation is provided in Section 2.3.
2.2.2.1 Pressurization system
In mono and bi-propellant propulsion options, a pressurization system is used for controlling the propellant pressure in the tank and for feeding it to the thruster combustion chamber. Nitrogen or helium are used as pressurant. They are inert gases with a low molecular weight. Helium would provide the lightest system but its leakage is difficult to prevent, therefore nitrogen is more often used [17]. There are two types of pressurization system for spacecraft application: blowdown and regulated option.
1. Blowdown is the lightest, simplest and hence more reliable type. On the other
(a) Mono-propellant. (b) Bi-propellant.
Figure 2.6: General schematic of liquid systems.
hand, the pressure at the thruster inlet will decrease during orbit operation due to fuel consumption. Initially the tank is loaded with propellant plus a certain quantity of pressurant gas. As the fuel is used, the fixed mass of gas expands in the ullage volume and the propellant pressure decreases. As a result, the mass flow rate at the thruster inlet decreases and the produced thrust is reduced. Mono-propellant systems can work properly even if sub- jected to the pressure decrease and nowadays they exclusively use blowdown systems. A simplified schematic of a blowdown system is shown in Figure 2.7(a).
2. Regulated system includes a pressure regulator between the pressurant gas tank and the propellant tank. The regulator continuously adjusts the pres- sure in the propellant tank, keeping it constant at the specific level for op- timal thruster performances. This type of pressurization system is essential for bi-propellant engines that must be fed with propellants at a constant flow
rate for maintaining constant the mixture ratio between fuel and oxidizer.
Simplified schematic of a regulated system is shown in Figure 2.7(b).
(a) Blowdown pressurization system (b) Regulated pressurization system
Figure 2.7: Pressurization system options, from [5].
2.2.3 Electric propulsion system
Electric propulsion (EP) systems use the electrical power to accelerate the working fluid (Xe is the most common medium) and generate thrust. There is not a theo- retical limit to the maximum velocity achievable by the exhaust gas however the power required may become too high for the spacecraft power generator capabil- ity. Moreover, the weight of the power subsystem increases with increasing power requirements. Nowadays the most promising devices in the domain of electric propulsion are those belonging to the electro-static category. Ion thrusters such as the Hall Effect Thruster (HET) are for instance included. In a ion thruster a high molecular weight gas is ionized and the resulting ions are accelerated through a strong static electric field. Electrons are then injected in the plume to neutralize it and avoid charging effects on the spacecraft, see Figure 2.8. Typical performance parameters are Isp=3000 s and T =0.001-0.2 N. The general schematic of the pro- pellant feed lines of an electric propulsion system is as for the cold gas system, seen in Figure 2.5.
2.3 Typical components
This section provides description of both components that are included in the gen- eral schematics showed in Section 2.2 and others that are used as ground support
Figure 2.8: Schematic of ion thruster, from [10].
equipment during tests.
2.3.1 Valves
Many types of valves are used in propulsion systems, some of which can be used for different applications. For instance, a hand valve (HV) can be used for isolation purposes and be called ”Manual Isolation Valve (MIV)” or alternatively for filling and draining a tank and be called ”Fill and Drain Valve (FDV)”. The division by type and application is summarized in Table 2.2. A brief description of each valve follows.
Table 2.2: Valve types and functions.
Type Application acronym
Hand Valve Isolation2(MIV), Service (FDV, FVV, TP), Flow control (NV)
Pyro-valve Isolation2
Solenoid Valve Flow control (FCV) Latch valve -
Check valve -
• Hand Valve (HV): manually operated valve actuated by rotation of a knob, see example in Figure 2.9. The general name hand valve is normally substi- tuted by a more specific one deriving from the valve application.
2Isolation valve (IV): when shut, it shuts-off a part of the propulsion system lines. It is used to isolate the propellant lines from the engine and the high pressure pressurant gas from the propellant tanks until the system is brought online. Moreover, it provides isolation of the tanks from the rest of the circuits during ground testing of the lines.
a) Manual Isolation Valve (MIV): it is mechanically locked in open position before launch.
b) Fill and Drain Valve (FDV): used on ground for system level testing and AIV activities, prior to launch for loading or draining propellant.
c) Fill and Vent Valve (FVV): used on ground for system level testing and AIV activities, prior to launch for loading or venting pressurant gas or for pressurizing the system.
d) Test Port Valve (TP): used as test port for leak checking during on- ground assembly and testing.
e) Needle Valve (NV): used for precise flow regulation during on ground activities. It has a narrow orifice with a long tapered seat and a needle- shaped plunger on the end of a screw, which fits the seat exactly. Turning the screw moves the plunger, opening or alternatively closing the flow passage. See example in Figure 2.10.
Figure 2.9: Hand Valve example, from [36].
• Pyro-Valve (PV): it is a one shot device that upon activation provides per- manent opening or closing of a line. When the triggering signal is sent to the valve a pyro cartridge is fired creating gas that expands into a confined volume moving a piston inside the valve body. The piston movement will open the flow path if the valve is NC, close the flow path if the valve is NO (in this case the PV belongs to the isolation category).
• Solenoid Valve (SV): it is an electrically operated, automatic valve. An electromagnetic coil is energized or de-energized to either open or close the
Figure 2.10: Needle Valve example, from [35].
valve. When energized, a magnetic field is generated and an internal plunger is moved against the action of a spring. When de-energized, the plunger returns to its original position by the spring action. The plunger position determines whether the flow passage is open or closed, see Figure 2.11. The coil must be continuously powered to maintain the magnetic field that holds the plunger in the commanded state, leading to a high electrical power con- sumption. Solenoid valves are used to control the flow of propellant at the thruster inlet and are referred to as Flow Control Valve (FCV). Both NC and NO solenoid valves exist.
Figure 2.11: Normally Closed solenoid valve, from [33].
• Latch Valve (LV): it is an electrically actuated, automatic valve that exploits a permanent magnet to keep the valve in the last commanded position (open or close) without being continuously energized, thus offering significant en- ergy saving compared to standard solenoid valves. When de-energized, the small magnet’s field is not enough to move the plunger. As soon as the valve is energized the magnetic field is intense enough to move the plunger and keep the valve in the commanded position. Valve’s position will change when an- other electrical input, of opposite polarity, is applied. Latch valves are used when the valve position will be kept for long periods of time, for instance it can be included as redundant isolation valve for specific line sections that could require isolation at any time during the mission. As an example, in Fig- ure 2.6(a) a LV is installed between the thrusters and the propellant tanks.
A failure of the thrusters would require their isolation from the propellant line, that is provided by energizing the NO LV.
• Check Valve (CV): ensures that fluids can flow only in the desired direction.
CVs are always included in the pressurant lines between the tanks of the fuel and oxidizer, to prevent their mixing inside the tubing, see Figure 2.6(b).
Example of check valve is shown in Figure 2.12.
Figure 2.12: Check Valve example, from [37].
2.3.2 Pressure Transducer
A Pressure Transducer (PT) converts the pressure signals to electrical signals which can be red by the on-board acquisition software (or a ground DAQ for test- ing). Pressure transducer only monitors the pressure without actively control it.
They can exploit different sensing principle, being differentiated in piezo-resistive3
3Piezo-resistive based transducers rely on the piezo-resistive effect which occurs when the electrical resistance of a material changes in response to applied mechanical strain. An electrical circuit then converts the change in resistance into a voltage output which is proportional to the applied pressure.
and piezoelectric4 sensors. Based on the reference pressure to which the measure- ment is compared, pressure transducers are distinguished in:
a) Absolute PT: it measures the pressure referring to a high vacuum reference, sealed behind their sensing diaphragm. Using an absolute PT eliminates error sources as thermal errors which would occur because of the thermal expansion-contraction of the gas sensed by the reference port. The working principle is shown in Figure 2.13.
Figure 2.13: Absolute pressure transducer, from [16].
b) Gauge PT: it measures the pressure relative to the atmospheric sea level ref- erence, that is 1013.25 mbar on average. The output of this type of sensor will change depending on the actual atmospheric condition, therefore its measure- ment is more prone to error. Gauge pressures higher and lower than ambient pressure are respectively measured as positive and negative pressures. The working principle is shown in Figure 2.14.
c) Differential PT: it measures the pressure difference between two pressure ports, each on one side of the sensing diaphragm. One should note that also the absolute and the gauge transducers are in reality differential sensors with respect to vacuum and atmospheric pressures, but usually the term differential is referred to the type of application where the two pressures on the diaphragm sides are ”process pressures”. Basing on the latter pressures,
Silicon is often used as piezo-resistive material [1]. This sensing technology is the most commonly used for general purpose pressure measurement.
4Piezoelectric based transducers rely on the piezoelectric effect, which occurs when a crystal reori- ents under stress forming an internal polarization. This polarization results in the generation of charge (therefore electrical potential) on the crystal face that is proportional to the applied stress. Quartz is mostly used as sensing material [1]. Piezoelectric transducers are the most suitable for measuring dynamic pressures.
Figure 2.14: Gauge pressure transducer, from [16].
differential pressure transducers can in turn be distinguished in wet/wet, wet/dry and dry/dry as per below:
• wet/wet: device which is compatible with liquid on both the positive and negative side process connections. A single diaphragm with fluid pushing on both sides or two separate sensors can be used to calculate the differential pressure.
• wet/dry: device which is compatible with a liquid on the positive pres- sure side and a dry/non-condensing gas on the negative pressure side.
This is most commonly found with compact designs where a gauge ref- erence design has been adapted to measure differential pressure;
• dry/dry: device which is only compatible with dry/non-condensing gas on both the positive and negative pressure side. Usually designed as single diaphragm where the diaphragm deflects in one direction or the other.
The working principle is shown in Figure 2.15.
2.3.2.1 Pressure Regulator
Pressure Regulator (PR) is used to control the fluid pressure in the lines. It reduces the inlet pressure to a desired lower value and keeps it constant, despite possible fluctuations at the inlet. Pressure regulator is constituted of three functional elements:
1. Restricting element: it is used to reduce the pressure, usually a spring loaded poppet valve.
Figure 2.15: Differential pressure transducer, from [16].
2. Sensing element: it senses the pressure changes in the system and drives the loading and restricting element to open or close the flow passage. Most commonly a diaphragm or a piston is used.
3. Loading element: it applies the force needed to the restricting element. Often it is a spring.
A general schematic of a pressure regulator is shown in Figure 2.16. At start, the spring keeps the pressure reducing valve open, allowing flow passage. As flow is established, the rise of the outlet pressure generates a force on the sensing element that counterbalance the spring force. The equilibrium between the two forces is reached at the set point of the pressure regulator. If the outlet pressure increases above the set point, the force generated by the sensing element will be greater than the force of the spring and the poppet will close the orifice.
Figure 2.16: Pressure regulator schematic, from [2].
2.3.2.2 Filter
Filter : they are used to prevent the contamination of the fluid flow by particles or debris that could damage sensitive components. Indeed, to achieve thrusters nominal operation a strict cleanliness requirement has to be met. See example in Figure 2.17.
Figure 2.17: Filter example, from [34].
Test Bench Requirements Definition
The test bench requirements have been derived starting from documents related to OHB ongoing projects. A test matrix containing the tests used as reference for the design of the bench has been created first. Then, the test requirement spec- ifications of every component belonging to the Euclid and Electra missions have been analyzed and the collected information have been used to draw a schematic for carrying each single different test and to list the set of requirements applicable to the test bench to be designed.
3.1 OHB references
The first step undertaken for the requirement definition of the test bench is to study the requirements specification of component belonging to some of the ongoing OHB projects that will be soon tested at qualification and acceptance stage. Even though only two projects are described in this document, note that the bench is designed to be used for testing of components of every future project as well.
3.1.1 Euclid
Euclid is a cosmology mission dedicated to studying the dark universe with un- precedented accuracy. Dark energy and dark matter are still mysterious elements of today’s “Standard Model’ of cosmology for which the majority of matter in
the universe is invisible (dark matter), and the universe is expanding at an in- creasing rate under the action of a still unknown energy source (dark energy).
Thanks to different number of investigation techniques in a very large survey over the full extragalactic sky, the Euclid payload will investigate the distance-redshift relationship and the evolution of the cosmic structures by measuring shapes and redshifts of distant galaxies. Euclid payload comprises a 1.2m diameter telescope and two scientific instruments: a visible-wavelength camera and a near-infrared camera/spectrometer [12].
The Euclid spacecraft is 4.5 meters long with a diameter of 3.1 meters and a mass of 2.1 tonnes. It will be operated from an orbit around the sun-earth L2 point located 1.5 million km from earth [13]. Euclid artistic representation is shown in Figure 3.1. For the Euclid mission, OHB Sweden has been appointed
Figure 3.1: Euclid artistic representation, from [11].
for the development of the Micro Propulsion Feed Assembly (MPFA) part of the Micro Propulsion System (MPS) and for the fluidic part of the Reaction Control System (RCS).
Euclid’s MPS is a cold gas propulsion system consisting of 6 pairs of 1-1000µN micro-thrusters capable of producing discrete variable thrust pulses. Within the MPS, the MPFA consists of the complete system but the thrusters and their elec- tronics. MPFA task is to feed the Micro Thruster Assembly (MTA) with pro- pellant, ensuring performances needed to meet the exceptional thrust accuracy,
resolution and stability required for fine attitude control and subsequent point- ing of the instruments on-board. The complete MPS is shown in Figure 3.2, the schematic of Euclid MPFA derives from the general one of cold gas system shown in Figure 2.5. Two thruster branches are included, a primary and a redundant one each equipped with 6 thrusters.
Figure 3.2: Euclid MPS, from [28].
Referring to Figure 2.5, Euclid lines upstream of the pressure regulators are HP, the lines downstream are indicated as VLP. The MPFA HP MEOP is 310 bar while the VLP MEOP is 4 bar. The expected maximum mass flow rate is 14 mg/s of N2. Note that the outlet pressure maintained by the pressure regulator during normal operation is 1.5 bar. The VLP MEOP is higher to keep into account possible pressure drifts caused by thermal effect or leakages through the regulator.
Euclid’s RCS is a Hydrazine mono-propellant propulsion system. Typical mono-propellant schematic is shown in Figure 2.6(a). Two thruster branches are used, a primary and a redundant one, each including 10 pairs of 20 N thrusters.
It is used for providing the DeltaV required for Transfer Correction Manoeuvres (TCM) to bring the spacecraft to the Lagrangian Point L2, and the periodic Station Keeping Manoeuvres (SKM). In addition, the RCS provides the torques necessary for stabilizing the spacecraft attitude after separation from the launcher, for man- aging the safe mode and for reaction wheels offloading [30]. The RCS lines are referred to as LP, their MEOP is 24 bar. The expected maximum mass flow rate is 108 g/s of Hydrazine.
3.1.2 Electra
The Electra mission belongs to the Electra Program, whose aim is to demonstrate the in-orbit capabilities of a fully electric propulsion (EP) geostationary platform.
To date, electric propulsion systems have only been used in research satellites or for orbit maintenance of telecommunications satellites. The concept of fully EP spacecraft has been already successfully tested by the Artemis satellite, operated by ESA, and the results showed that electric propulsion systems reduce propel- lant mass requirements by up to 90% compared with chemical propulsion systems.
This allows for a launch mass reduction of almost 50%. The mass saving has as an advantage the possibility to use smaller launch vehicles while carrying payload ca- pabilities equivalent to larger chemical propulsion mid-sized satellites, translating into significant economic benefits.
OHB Sweden is in charge for the development of the Electric Propulsion Sub- system (EPPS) and for the Attitude and Orbit Control System (AOCS) of the Electra spacecraft. The latter will not be discussed here.
Figure 3.3: Electra artistic representation.
The EPPS relies on four HET thrusters fed by Xenon gas, two nominal and two redundant, mounted on two articulated booms as shown in Figure 3.4.
Figure 3.4: Electra booms configuration, from [31].
The architecture of the Xenon feed lines is derived from the general schematic
of cold gas system shown in Figure 2.5. Referring to Figure 2.5, the lines upstream the pressure regulator are HP, the lines downstream are LP. The HP MEOP is 186 bar while the LP MEOP is 15 bar. The expected maximum mass flow rate is 250 mg/s of Xe. The pressure regulator will reduce the pressure from 186 to 2.55 bar. Note that the nominal LP maintained by the regulator is different from the LP MEOP of 15 bar. The LP MEOP is set at 15 to keep into account pressure increase given by thermal effects and leakages across the regulator.
Each EP boom is driven by three actuators, A1,A2,A3 in Figure 3.5, for direct- ing the thrust vector in the required orientation. The EPPS is in charge for the orbit raising to GEO, for final orbit acquisition and station keeping as well for the disposal to a graveyard orbit at EOL. Additionally, the EPPS will be used for all the momentum management operations. A cold gas system, not presented here, is used to provide the torques necessary for spacecraft de-tumbling after separation from the launcher and for attitude control during the safe mode.
Figure 3.5: Electra booms kinematic, from [31].
3.2 Test matrix
To define a test matrix, the first step is to identify the tests that have to be carried for each component and to define the relevant fluid medium. This has been done by reviewing the existing requirements specification documents for each component of each project. Table 3.1 includes all the fluidic tests that shall be carried on components at different verification stages. Table 3.2 summarize the tests that have been considered for the design of the test bench. Filter and MIV acceptance testing for the Euclid and Electra projects will be carried out soon, therefore
the corresponding tests are prioritized. Service valve (SV), pyro-valve (PV) and tank are not considered since they are not tested in OHB propulsion laboratory.
Moreover, burst pressure is not considered for safety reasons and filtration level is excluded because it would require the insertion of dust particles in the flow that would contaminate all the lines of the bench.
Table 3.1: Fluidic tests.
*Included in the acceptance test block for the Euclid pressure regulators.
Test type Qualification Acceptance
Internal leak X X
External leak X X
Reverse internal leak X X
Flow rate X X
Pressure drop X X
Water hammer X
Slam start X
Noise on regulated pressure X [X]*
Proof pressure X X
Burst pressure X
Pressure cycling X X
Bubble point X X
Filtration capacity X
3.3 Test description & requirements
The schematic and requirements for each test have been extrapolated from tests that have already been carried out at OHB Sweden and from the requirement specification of each component of each project. They are presented in the next sections, together with a brief explanation of the procedure to be followed.
3.3.1 Leak tests
Leak is a flow of medium through the wall or the seat of a component. No com- ponent can be absolutely leak-tight and it does not need to be. The essential is that the leak rate must be low enough that the required operating pressure is not influenced. To calculate the leak rate, the pressures acting on either side of the
Table3.2:Testmatrix,thetestmediumisindicatedinsidebrackets. *Theuseofpropellantsimulantscaninfuturebeconsidered.Ethanolissuggestedin[26]asthebesthydrazinesimulant, HFE7100isrecommendedin[6]forsimulatingNTO. TEST Componentinternal leakexternal leakreversed int.leakflowratepressure dropwater hammerslam startnoise onreg. pressure
proof pressurepressure cyclingbubble p MIVx (He)x (He)x (He)x (Xe,N2)x (Xe,N2)x (N2)x (N2,He)x (N2,He) LVx (He)x (He)x (He)x (Xe,N2, water*)
x (Xe,N2, water*)
x (water*)x (N2,He)x (N2,He) SVx (He)x (He)x (He)x (Xe,N2)x (Xe,N2)x (water*)x (N2,He)x (N2,He) CVx (He)x (He)x (Xe,N2)x (Xe,N2)x (N2,He)x (N2,He) PTx (He)x (water*)x (N2)x (N2,He) PRx (He)x (He)x (Xe,N2)x (N2)x (N2,Xe)x (N2,He) Filterx (He)x (Xe,N2, water*)
x (Xe,N2, water*) x (N2,He)x (N2,He)x (IP Orificex (Xe,N2)x (Xe,N2)
wall/seat and the nature of the test medium (viscosity, molar mass) shall be de- fined. Helium is mostly used as tracer gas as it is one of the smallest gas molecules, it is inert that means it will not interact with any of the materials within the part to be tested and it is quite economic. A mass spectrometer tuned to detect helium is used to evaluate the leak rate ensuring that only helium and no other gases can influence the test. If a different gas is used or if the requirements specification states the maximum leak rate referred to a different gas than helium, a conversion formula shall be applied to find the equivalent leak rate. In this regard, it must be noted that different formulas apply to laminar viscous and molecular flow. The boundary between these areas is very difficult to determine but, as a guideline, the following can be assumed:
Ql=
< 10−7 mbar·l/s ⇒ molecular flow
10−1÷ 10−6 mbar·l/s ⇒ laminar viscous flow
(3.1)
Molecular flow exists in small leaks and at low pressures. At molecular flow each molecule travels independently of other molecules. The mean free path is larger than the diameter of the leak capillary. Therefore it is possible, that a single molecule travels against the general flow direction, because molecules do not collide with each other, but only with the walls as seen in figure 3.6. The leak rate at molecular flow is calculated using the formula of Knudsen as follow:
Ql =
√2π 6
r R · T mm
d3
l (P1− P2) (3.2)
where
R = universal gas constant;
T = absolute temperature (K);
mm = molar mass;
P1 = higher pressure;
P2 = lower pressure;
d = diameter of the leak;
l = length of the leak.
Assuming that the geometry dimensions of the leak and the pressure conditions do not change during the period of measurement and considering two types of
gases, equation 3.2 implies that the leak rate in a molecular flow follows inverse proportion to the square root of the relative molar mass of the gases:
Qa·√
mma = Qb ·√
mmb. (3.3)
Figure 3.6: Molecular flow path of a single molecule, from [22].
Laminar viscous flow is defined as a parallel flow of molecules in a pipe with parabolic velocity distribution in the transverse direction. Figure 3.7 shows the flow path for a laminar flow inside a pipe. Poiseuille law is used to calculate the
Figure 3.7: Laminar flow inside a pipe. The maximum velocity is in the centerline of the pipe and decreases towards the walls, from [22].
leak rate for laminar conditions as follow:
Ql = π · r4
16 · η · l(P12− P22) (3.4) where
r = radius of the pipe;
η = dynamic viscosity of the gas [bar·s];
l = length of the pipe;
P1 = higher pressure;
P2 = lower pressure.
Assuming that the geometry dimensions of a leak do not change during the period
of measurement and under the same pressure conditions, from equation 3.4 it can be deduced that when the type of gas changes, the leak rate follows an inverse proportion to the dynamic viscosities of the gases:
Qa· ηa= Qb· ηb. (3.5)
Propulsion system components are leak tested with respect to internal, external and reverse internal leak. Test specifications and schematic for each of them are presented in the next subsections.
3.3.1.1 Internal leak
Internal leak consists of leakage across the valve or regulator seat when in closed position. Being the leak only across an internal component of the item, the leaked fluid remains contained within the piping system. The goal of the test is to evaluate the magnitude of internal leak across the component and verify that its rate does not exceed the level specified in the requirement. For this test, Helium is used as test medium.
Simplified schematics for valve and regulator testing are shown in Figure 3.8.
For the valve test, a Helium detector is connected downstream. First, the valve is set to the closed position, and the He detector is used to vacuum the line down- stream the valve. Then pressure is applied upstream until reaching the desired level and the amount of time specified by the requirement is waited. The possible Helium leak across the valve will be detected by the He detector. For the regulator test, a high resolution, high accuracy pressure transducer is used to measure the leakage rate instead of the He detector. First, inlet pressure is supplied until the regulator goes into lock-up condition1 and then increased until the regulator’s inlet MEOP. The possible leakage across the regulator will result in a pressure increase downstream, recorded by the PT.
1Lock-up condition: condition in which the regulator is completely shut off not allowing flow.
(a) Valve test.
(b) Pressure regulator test.
Figure 3.8: General schematic of the internal leak test.
3.3.1.2 External leak
External leak consists of leakage from inside the item out to the ambient. The goal of the test is to evaluate the leak when the item is pressurized at different level up until its MEOP for a defined amount of time and verify that the leakage rate does not exceed the maximum specified in the requirement.
Simplified schematics for external leak testing of the different items are shown in Figure 3.9. The tested item is placed inside a vacuum chamber to which the He detector is connected. Then the procedure to be followed is different between the pressure regulator and all the other items. Testing the pressure regulator requires the possibility of pressurizing both the upstream and downstream lines. In fact, the pressure regulator external leakage shall be measured with the inlet and outlet lines pressurized at their respective MEOPs with respect to ambient, that are higher than the nominal working pressures. First, vacuum is created into the vacuum chamber and pressure is supplied at the regulator’s inlet until it reaches the lock- up condition. Then its upstream and downstream lines are pressurized to their respective MEOPs. Pressures and leakage rate are continuously monitored and registered by the pressure transducers and the He detector for the entire duration
of the test. For all the other items, vacuum is created inside the chamber and inlet pressure is supplied to the component until reaching the desired level. Note that valves are tested in closed position. An end-line cap is added at the tested item outlet to avoid contributes given by the internal leak. The supplied pressure and the leakage rate are continuously monitored and registered.
(a) Valve, PT, filter test.
(b) Pressure regulator test.
Figure 3.9: General schematic of the external leak test.
3.3.1.3 Reverse Internal leak
The reverse internal leak consists of internal leak when pressure is applied to the item in the outlet to inlet direction. The test concept, schematic and procedure is the same as for the internal leak test described in Section 3.3.1.1, the only difference being the pressurization direction.
3.3.2 Flow rate & Pressure drop
The goal of this test is to verify that the item under test is able to deliver a certain mass flow rate with a maximum pressure drop from inlet to outlet not exceeding a certain value. The achievable flow rate and corresponding pressure drop are evaluated for different inlet pressures up to the item’s MEOP. The required mass flow rate and the maximum allowable pressure drop are specified in the requirement specification of each item. Note that the flow rate test of the pressure regulator is used to evaluate the stability of the regulated pressure and the extent of the Joule-Thomson effect, indeed pressure drop in pressure regulator has no meaning.
The test medium comprehends gaseous nitrogen, xenon or water, depending on the project and on the propulsion system type to which the tested item belongs to.
Simplified schematics are shown in Figure 3.10 and 3.11. The differential pres- sure transducer measures the pressure drop across the tested item and a mass flow meter is used to measure the flow, that is adjusted to the desired level by manually opening/closing a downstream needle valve. For the pressure regulator test, the item is placed in the vacuum chamber and a pressure transducer in the low pressure line is used to monitor the regulated pressure stability. It shall be noted that the simplified schematics do not include possible items needed for ther- mal control of the flow, that could be required depending on the magnitude of the Joule Thomson (JT) effect. Together with the JT effect, the choked flow condition shall also be taken into account, as explained in the next sections.
3.3.2.1 Joule Thomson effect
The Joule–Thomson effect describes the temperature change of a real fluid when subjected to a sudden pressure change. For instance, the temperature change per- taining to the Joule-Thomson effect occurs in pressure regulator where the fluid
(a) Valves, filter, orifice test using gas medium.
(b) Valves, filter, orifice test using liquid medium.
Figure 3.10: General schematic of the mass flow and pressure drop test.
at pressure P1 flows into a region of lower pressure P2, following an irreversible expansion. It can be proved mathematically that during this expansion, enthalpy remains unchanged (isenthalpic process) while internal energy does not thus re- sulting in a temperature change.
While exploited in refrigeration and air conditioner systems, the Joule-Thomson effect is not necessarily desirable in propellant pressure regulation. Indeed, an ex- cessive cooling or heating of the propellant can lead to malfunctioning or failure of the system. The extent of the temperature change depends on the initial and final state and on the type of fluid and must be evaluated case by case. If needed, heat- ing or cooling elements shall be used to compensate for the temperature change.
To calculate the magnitude of the needed heating or cooling, the process is re- garded as non-isenthaplic and the change of enthalpy between the initial (Pi, Ti) and final (Pf, Tf) states is multiplied by the mass flow rate, as shown in Table 3.3, an example taken from the Electra MPR phase B1 testing [18].
Figure 3.11: Pressure regulator test using gas medium.
Table 3.3: Joule-Thomson Assessment, Electra MPR testing phase B1
Test Medium Xenon
Mass flow rate m = 0.2 g/s˙ Initial State
Ti = 50◦C Pi = 186 bar hi = 61.467 J/g Final State
Tf = 20◦C Pf = 1 bar hf = 116.3 kJ/kg
∆h = 54.833 J/g Q = ˙m · ∆h= 11.8 W
3.3.2.2 Choked flow condition
Choked flow arises in real fluids where compressibility effects play a major role.
Choked condition refers to the fact that the fluid velocity can no increase further than the local speed of sound and therefore, at fixed upstream conditions, also the mass flow can not increase.
The choked flow condition is at M =1 and is obtained for a specific isentropic pressure ratio calculated using:
P0 P =
1 + γ − 1 2 M2
γ−1γ
→ with M=1 → P0
P = γ + 1 2
γ−1γ
. (3.6)
In other words, for each pressure ratio greater or equal to the one calculated in equation 3.6 the flow is choked.
Starting from the mass flow conservation ˙m = ρ·u·A, using isentropic flow relations and the equation of state, the compressible form of the mass flow rate equation can be derived:
˙
m = AP0 r γ
RT0M
1 + γ − 1 2 M2
− γ+1
2(γ−1)
. (3.7)
The mass flow rate at choked condition can be calculated substituting M =1 into equation 3.7, giving:
˙
m = AP0 r γ
RT0
2
γ + 1
2(γ−1)γ+1
. (3.8)
Note that in all previous equations A is the flow passage area;
P0 is the absolute upstream pressure;
P is the absolute downstream pressure;
γ = CCp
v is the heat capacity ratio;
R = Cp− Cv is the specific gas constant;
T0 is the absolute upstream temperature;
M is the local speed of sound.
It must be noted that even though the fluid velocity will remain constant also for decreasing downstream pressure, the mass flow rate can be increased increasing the upstream pressure, which will in turn increase the upstream density of the fluid.
As an example, the choked flow condition calculations from the Electra MPR phase B1 testing are presented in Table 3.4. Note that during this test Xenon is released in ambient air.
3.3.3 Water hammer
Water hammer consists in a pressure peak generated inside a piping system and in the corresponding shock wave that is transmitted throughout the fluid. Water hammer is caused when a fluid in motion is forced to stop suddenly: the momentum of the fluid traveling in its forward direction will force the fluid to keep moving in that direction but as the fluid stops, its rapid change of momentum generates a
