design and implementation of ground support equipment for ......ekinetic – kinetic energy of gtv...
TRANSCRIPT
Design and Implementation of Ground Support Equipment for
Characterizing the Performance of XPOD and CNAPS
&
Thermal Analysis of CNAPS Pressure Regulator Valve
by
Mohamed R. Ali
A thesis submitted in conformity with the requirements
for the degree of Master of Applied Science
Graduate Department of Aerospace Engineering
University of Toronto
© Copyright by Mohamed R. Ali 2009
ii
Abstract
Title: Design and Implementation of Ground Support Equipment for
Characterizing Performance of XPOD and CNAPS & Thermal Analysis of
CNAPS Pressure Regulator Valve
Degree: Master of Applied Science
Conferred: 2009
Name: Mohamed R. Ali
Department: Aerospace Engineering
University: University of Toronto
As the potential uses of nanosatellites become apparent, their numbers keep increasing.
This is evident at the Space Flight Laboratory (SFL) located at the University of Toronto
Institute for Aerospace Studies which has seen a rapid growth in satellite missions in
recent years. By leveraging the use of the Generic Nanosatellite Bus developed at SFL,
satellites can be rapidly developed to carry science payloads or demonstrate the
capabilities of new technologies on orbit.
Testing satellite systems in an Earth environment is an important step in qualifying them
for space. This thesis describes the development of ground support equipment for testing
SFL’s nanosatellite separation system, XPOD, and characterizing the performance of the
propulsion system, CNAPS. Also, the thermal behaviour of a pressure regulator valve on
CNAPS is examined for various flow conditions.
iii
Acknowledgements
The opportunity to work on satellites and space hardware as part of my thesis work has
been a tremendous experience from the very beginning. The exposure to these very real
missions and their respective schedules has allowed me to gain valuable knowledge and
insight into the life cycle of space missions. For this, I am grateful to Dr. Robert Zee who
gave me the opportunity to be a part of SFL. His vision for providing low cost access to
space has, and will continue to revolutionize the Canadian space industry.
Two people I worked closely with during my graduate work at SFL are Stephen Mauthe
and Freddy Pranajaya. Their feedback and assistance has been very valuable in
completing this thesis. Daniel Kekez has been helpful in all respects, be it explaining the
basics of breadboard circuits, or answering questions related to all things electrical. CanX
program manager, Cordell Grant, was always willing to help whenever I approached him
with design related queries. All SFL staff members have provided me with guidance at
some point in time, and I am thankful to them.
Past and present students at SFL have made some days more bearable than they would
have been. I am appreciative of their company.
My family deserves my utmost gratitude and respect for their unfaltering support
throughout my years at university.
iv
Table of Contents
INTRODUCTION ........................................................................................................................................ 1
1. CANX PROGRAM............................................................................................................................. 2
2. MISSION TIMELINE........................................................................................................................ 3
2.1. PAST M ISSIONS............................................................................................................................ 3
2.2. FUTURE M ISSIONS........................................................................................................................ 4
3. XPOD SEPARATION SYSTEM....................................................................................................... 5
4. CNAPS ................................................................................................................................................. 7
5. THESIS OUTLINE............................................................................................................................. 8
CHAPTER 1 XPOD GROUND SUPPORT EQUIPMENT
1. INTRODUCTION............................................................................................................................... 9
2. LITERATURE REVIEW................................................................................................................. 11
3. REQUIREMENTS............................................................................................................................ 12
4. DESIGN ............................................................................................................................................. 15
4.1. GROUND TEST VEHICLE (GTV) .................................................................................................16
4.1.1. Supporting Structure............................................................................................................. 16
4.1.2. Satellite Interface Adapter.................................................................................................... 18
4.2. REDUCED FRICTION CONDITION................................................................................................. 22
4.2.1. Criteria Weighting and Option Rating ................................................................................. 24
4.3. XPOD MOUNT........................................................................................................................... 27
4.4. MASS MODEL............................................................................................................................. 29
4.5. GNB SPARE STRUCTURE – SATELLITE MODEL.......................................................................... 30
5. MASS BUDGET................................................................................................................................ 30
6. SAFETY AND RELIABILITY OF GTV ....................................................................................... 31
7. DATA COLLECTION METHODS................................................................................................ 34
8. DEPLOYMENT ANALYSIS........................................................................................................... 34
9. TESTING........................................................................................................................................... 38
10. CONCLUSION............................................................................................................................. 41
v
CHAPTER 2 THRUST STAND DESIGN AND CALIBRATION
1. INTRODUCTION............................................................................................................................. 42
2. LITERATURE REVIEW................................................................................................................. 43
3. REQUIREMENTS............................................................................................................................ 44
4. DESIGN ............................................................................................................................................. 45
5. DATA COLLECTION ..................................................................................................................... 53
6. ANALYSIS & CALCULATIONS................................................................................................... 55
6.1. THRUST COMPONENTS............................................................................................................... 60
7. CALIBRATION................................................................................................................................ 60
7.1. SETUP......................................................................................................................................... 60
7.2. DATA COLLECTION .................................................................................................................... 61
7.3. RESULTS.................................................................................................................................... 63
8. RECOMMENDATIONS AND FUTURE WORK......................................................................... 65
CHAPTER 3 CNAPS PRESSURE REGULATOR VALVE THERMAL ANALYSIS AND TESTING
1. INTRODUCTION............................................................................................................................. 66
2. DESIGN ............................................................................................................................................. 67
3. ANALYSIS ........................................................................................................................................ 68
3.1. RELEVANT THEORY ................................................................................................................... 69
3.1.1. Isentropic Flow..................................................................................................................... 69
3.1.2. Thermodynamics................................................................................................................... 70
3.1.3. Heat Transfer........................................................................................................................ 72
3.2. THERMAL MODELLING............................................................................................................... 75
4. BENCH TESTING............................................................................................................................ 79
4.1. SETUP......................................................................................................................................... 79
4.2. DATA COLLECTION..................................................................................................................... 80
4.3. TESTS......................................................................................................................................... 83
5. CONCLUSION ................................................................................................................................. 85
CONCLUSION........................................................................................................................................... 86
REFERENCES ........................................................................................................................................... 88
vi
List of Figures
Figure 1 GNB architecture.................................................................................................. 3
Figure 2 CanX-3 (left) and CanX-4 (right) with payloads ................................................. 5
Figure 3 XPOD family of separation systems .................................................................... 6
Figure 4 CNAPS internal layout ......................................................................................... 8
Figure 5 XPOD GNB with satellite .................................................................................... 9
Figure 6 Previous setup for testing XPOD deployment.................................................... 10
Figure 7 Supporting structure of GTV.............................................................................. 17
Figure 8 Exposed faces of satellite when integrated with XPOD..................................... 17
Figure 9 Shims used to adjust height of base block.......................................................... 18
Figure 10 Earlier design of H-frame, allowing satellite to rotate independent of GTV ... 20
Figure 11 Method of fastening H-frame to base block in earlier design .......................... 21
Figure 12 Components of H-frame................................................................................... 21
Figure 13 Method of attaching H-frame to satellite.......................................................... 22
Figure 14 Tipping analysis for XPOD fastened to Mount................................................ 28
Figure 15 Mass model halves showing optional masses inside........................................ 29
Figure 16 Mass models of components inside GNB tray................................................. 30
Figure 17 Tipping analysis of satellite and GTV.............................................................. 33
Figure 18 Accelerometer used for deployment tests (courtesy SparkFun Electronics).... 34
Figure 19 Estimated deployment speeds of satellite based on analytical model .............. 36
Figure 20 Setup of GSE for deployment tests .................................................................. 38
Figure 21 Sample plot of data collected by accelerometer during test ............................. 39
Figure 22 Deployment speeds measured during tests....................................................... 39
Figure 23 Fulcrum for thrust stand provided by pivot blocks .......................................... 47
Figure 24 Schematic of thrust stand identifying forces and their locations...................... 48
Figure 25 Features of multiplier arm ................................................................................ 49
Figure 26 Method of fastening CNAPS to thrust stand.................................................... 50
Figure 27 H-frame used for measuring misalignment of thrust vector............................. 51
Figure 28 CNAPS configuration for measuring thrust vector misalignment ................... 51
vii
Figure 29 Adding counterweights to multiplier arm......................................................... 52
Figure 30 Thrust stand with CNAPS attached to it........................................................... 53
Figure 31 CNAPS orientation for Case 1 ......................................................................... 56
Figure 32 Nozzle placement with respect to centre of CNAPS........................................ 57
Figure 33 CNAPS orientation for Case 2 ......................................................................... 58
Figure 34 CNAPS orientation for Case 3 ......................................................................... 59
Figure 35 Setup of thrust stand inside SFL vacuum chamber .......................................... 61
Figure 36 Sample plot of measured force showing data extraction range........................ 62
Figure 37 Sample plot of measured force for small mass and scaling factor ................... 63
Figure 38 Plot showing relationship between measurement error and scaling factor ...... 64
Figure 39 Regulator valve fastened to CNAPS structure .................................................68
Figure 40 Regulator valve cross section (Courtesy Mike Borla, The Lee Co.)................ 68
Figure 41 Equivalent circuit diagram for thermal resistsance ..........................................73
Figure 42 Control volume around regulator valve............................................................ 73
Figure 43 Temperature profile of valve without gap filler ............................................... 76
Figure 44 Temperature profile of RV for nominal operations with gap filler in use........ 76
Figure 45 Temperature profile of RV for extended operations with gap filler in use ...... 77
Figure 46 Temperature profile of test RV with gap filler in use ...................................... 78
Figure 47 Setup for thermal testing of RV........................................................................ 79
Figure 48 Test valve fastened to CNAPS with Delrin block (front view)........................ 80
Figure 49 Schematic of breadboard circuit ....................................................................... 82
Figure 50 Test valve with regions of interest identified ................................................... 83
Figure 51 Minimum temperatures measured at RV exit port for different on-times........ 84
viii
List of Tables
TABLE 1 LIST OF CRITERIA USED TO EVALUATE OPTIONS FOR REDUCING FRICTION
EFFECTS ..................................................................................................................................................... 24
TABLE 2 MASS BUDGET OF GTV.......................................................................................................... 31
TABLE 3 SIZING OF MULTIPLIER ARM BASED ON FORCE SENSOR LIMITS .............................. 48
TABLE 4 PREDICTED AND MEASURED MINIMUM TEMPERATURES AT VALVE EXIT PORT. 84
Acronyms
AIS Automatic Identification System BRITE BRIght Target Explorer CanX Canadian Advanced Nanospace eXperiment CN CNAPS CNAPS Canadian Nanosatellite Advanced Propulsion System COTS Commercial Off-The-Shelf DAQ Data Acquisition Unit ESC Electro-Static Combs GNB Generic Nanosatellite Bus GSE Ground Support Equipment GTV Ground Test Vehicle LV Launch Vehicle MOST Microvariability and Oscillations of STars NANOPS NANOsatellite Propulsion System NTS Nanosatellite Tracking Ships RV Regulator Valve SFL Space Flight Laboratory T-POD Tokyo/Toronto Picosatellite Orbital Deployer TS Thrust Stand TVAC Thermal VACuum UHF Ultra High Frequency UTIAS University of Toronto Institute for Aerospace Studies VHF Very High Frequency XPOD eXperimental Push-Out Deployer
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List of Symbols
Chapter 1
g – gravitational acceleration
mplate – mass of XPOD Mount base plate rplate – distance of base plate CG from pivot
mXPOD – mass of XPOD rXPOD – distance of XPOD CG from pivot
msat – mass of satellite rsat – distance of satellite CG from pivot
Ekinetic – kinetic energy of GTV Epotential – potential energy of GTV
vCM – speed of GTV-satellite mass centre hmax – max. height above mass centre
Chapter 2
TF – CNAPS thrust in CN frame Tr – position of CNAPS wrt TS frame
SF – force measured by sensor, TS frame Sr – position of sensor wrt TS frame
TSℑ – thrust stand reference frame CNℑ – CNAPS reference frame
CNTSC , – transformation matrix from CN to
TS reference frame
OM – moment about TS reference frame
Chapter 3
mɺ – mass flow rate M – mach number
pO – stagnation pressure TO – stagnation temperature
A* – flow area corresponding to M = 1 R – gas constant
γ – specific heat ratio Eɺ – power
h - enthalpy ∆V – change in velocity
k – thermal conductivity l – dimension parallel to heat flow
A – cross sectional area, perpendicular to
heat flow
1
Introduction
Like computers from a generation ago, satellites were once considered to be big and
bulky objects, with appendages protruding from all sides. Less than half a century later,
technology has afforded the miniaturization of satellites to a level where one can fit into
the palm of an average human, weighing only a few grams. This reduction in size and
mass may be attributed to human ingenuity and creativity, as much as it can to
advancement in the applied science disciplines.
In the previous decade, smaller satellites have become more prevalent and the small
satellite community has grown as the capabilities of these satellites are realized. Smaller
satellites generally come with a small budget and schedule, and therefore provide the
added advantage of being used in educational institutions for providing hands-on
experience to students.
One very successful example of this is the Space Flight Laboratory, located at the
University of Toronto Institute for Aerospace Studies. Established in 1998 by Dr. Robert
Zee, SFL employs the ‘microspace’ philosophy in its approach to satellite engineering.
Small, discipline-oriented groups work in collaboration to integrate all subsystems into
one complete spacecraft. SFL contributed to Canada’s first space telescope, Micro-
variability and Oscillation of STars, the success of which gave birth to the laboratory’s
Canadian Advanced Nanosatellite eXperiment (CanX) program. Since then, SFL has
grown in size to include full-time staff members, many of them SFL-graduates, who
provide mentorship to students. The organization consists of two branches:
� The Satellite Systems Group is responsible for the design, assembly and operations of
all SFL built satellites.
2
� The Advanced Systems Group is responsible for special projects related to satellite
development, such as separation systems and mission specific payloads.
In April 2008, SFL achieved another milestone when it launched two more satellites into
orbit. CanX-2 and NTS (designated CanX-6) were launched atop the PSLV encased in
SFL’s very own separation systems. Another first for this launch campaign was the fact
that NTS was designed and built within seven months from the time the project was first
proposed, proving the effectiveness of utilizing the ‘microspace’ approach.
1. CanX Program
As a follow on to the successful MOST mission, the CanX program was founded in 2001
at SFL as a means to provide students with education and training in the design and
integration of a satellite. The program aims to give students the opportunity to be part of
the complete life-cycle of a satellite, from initial design, to assembly and testing, and
finally on-orbit operations. Graduate students work on particular satellite subsystems and
typically get to see their efforts launched into orbit as well, all in the time it takes them to
complete a master’s thesis.
To standardize the satellite design process, the generic nanosatellite bus was developed
by SFL staff and students. Having a cubic form factor, with sides measuring 20 cm, the
GNB aims to reduce recurring design and engineering costs for each subsystem. The
design consists of two trays that provide the primary structure for the satellite, to which
remaining modules and components are attached. Six panels cover the trays from all sides
to provide an enclosed spacecraft as shown in Figure 1. Space between both the trays is
allotted for payloads. The trays themselves contain subsystem components that support
the mission, such as attitude control, computer, and communications systems.
3
Figure 1 GNB architecture
2. Mission Timeline
2.1. Past Missions
MOST was the first satellite that was designed and assembled at SFL. Following that, the
CanX-2 mission was initiated, with the goal of being a technology demonstrator for
testing satellite technology. Some of the technology on board included UHF and S-band
radios, a reaction wheel, and a cold gas propulsion system. The experience gained from
testing these technologies on orbit would be applied to future design iterations. CanX-2
also carried onboard science experiments for Canadian scientists.
NTS was a joint mission between SFL and COM DEV for demonstrating the capability
of detecting automatic identification signals on orbit for tracking ships. The payload was
supplied by COM DEV, and was integrated with a GNB structure at SFL. MOST has
exceeded its mission life of one year and continues to produce useful data. CanX-2 and
4
NTS have celebrated their one-year anniversary and continue to exceed their one-year
mission life.
2.2. Future Missions
Upcoming missions at SFL have more challenging mission objectives to accomplish. The
technologies used on these missions are the next generation of those that were used on-
board CanX-2 and NTS. One of these is the BRITE mission, designated CanX-3, that
aims to examine the life cycle of stars by observing their relative intensity using precise
differential photometry [1]. Three reaction wheels provide three axis attitude control to
within one arc-minute. Attitude knowledge is provided by a star tracker. The BRITE
mission consists of a constellation of four satellites that will function in pairs. Each
satellite in a pair has a different optical filter on its telescope to detect different
wavelengths of light. Figure 2 shows the BRITE satellite’s layout.
Nanosatellites can also be used as a platform to test formation flying, a subject that is
heavily researched for space applications. The CanX-4/-5 mission developed at SFL aims
to do just that by demonstrating autonomous formation flying capability with
nanosatellites. The mission involves deploying two nanosatellites attached together,
separating them, and then maintaining them in formation with respect to each other. Both
satellites are identical in form and function, and include several new technologies such as
an inter-satellite separation system, a propulsion system, and an inter-satellite radio link.
The internal layout is shown in Figure 2. Manoeuvring ability will be provided by the on-
board liquid-fuel, cold gas propulsion system developed at SFL. The mission objective is
to achieve cm level knowledge in relative position and sub-metre accuracy in control [2].
5
Figure 2 CanX-3 (left) and CanX-4 (right) with payloads
AISSat-1 is a nanosatellite mission also based on the GNB architecture. The goals of this
mission are similar to those of the NTS mission: to demonstrate the reception of AIS
signals from space. A secondary objective of the mission is for the payload to triangulate
the position of vessels through continued observations of their transmitted signals [3].
AISSat-1 utilizes the GNB architecture with some modifications to accommodate the
payload antenna.
3. XPOD Separation System
SFL also develops its own line of nanosatellite separation systems, called the XPOD. The
XPOD is meant to provide a customizable ejection system for any size of nanosatellite.
The present design has evolved over many iterations and improvements to the previous
generation of XPOD’s. Yet, the actual mechanism that separates the satellite from the
launch vehicle has remained the same, thereby maintaining the space heritage achieved
by earlier XPOD launches.
Various models and sizes of the XPOD have been developed, targeting different classes
and form factors of satellites (Figure 3). The XPOD Single, and XPOD Triple can
accommodate satellites with a 10 x 10 cm cross section, adhering to the CalPoly CubeSat
6
standard. The more recently developed XPOD GNB and XPOD DUO are meant to
accommodate the GNB cross-section of 20 x 20 cm, for satellites weighing up to 7.5 kg
and 14 kg respectively. These two XPOD designs were based on a semi-enclosed concept
to accommodate fixed spacecraft appendages.
XPOD Single: 10x10x10 cm satellite
XPOD Triple: 10x10x30 cm satellite
XPOD GNB: 20x20x20 cm satellite
XPOD DUO: 20x20x40 cm satellite
Figure 3 XPOD family of separation systems
7
The XPOD functions as a jack-in-the-box separation system, with a main spring that
pushes the satellite out at the desired time. Prior to deployment, the satellite is secured
firmly within the XPOD, minimizing any movement, and the risk of damage. Each
XPOD undergoes rigorous testing to comply with multiple launch vehicle requirements.
The deployment process is characterized through ground deployment tests [4], [5].
4. CNAPS
The Canadian Nanosatellite Advanced Propulsion System (CNAPS) is the next
generation version of the Nanosatellite Propulsion System (NANOPS) that was
demonstrated on orbit aboard the CanX-2 satellite. While the primary purpose of
NANOPS was to demonstrate nanosatellite propulsion capability, CNAPS is to be used
on the CanX-4 and CanX-5 formation flying demonstration mission. The system
architecture and lessons learned from NANOPS were carried over to CNAPS, and hence
both propulsion systems share a fair amount of similarity.
CNAPS uses liquefied sulphur hexafluoride (SF6) as the propellant, which is stored on
board in liquid form. SF6 is regulated to a desired pressure and then vented out a four-
nozzle system to provide the propulsive force. Figure 4 shows the internal layout of
CNAPS. Precise control over the thrust magnitude is required to accomplish the
formation flying goals of the mission, and this is achieved by using pulse width
modulation techniques to control the pressure and flow rate of SF6 within CNAPS. The
cold-gas propulsion system has an Isp of 35 s and can produce a minimum thrust of 5 mN
from each thruster, to a resolution of 0.5 mN [6].
8
Figure 4 CNAPS internal layout
5. Thesis Outline
This thesis examines the characterization of the XPOD and CNAPS performance in
further detail. Chapter 1 details the design, development, and implementation of ground
support equipment relating to characterizing the XPOD deployment process. Results of
preliminary tests are summarized at the end of the chapter.
Chapter 2 deals with the design of a thrust stand for measuring the performance of
CNAPS. Details about the capabilities and calibration of the stand are mentioned, along
with a summary of the results.
The thermal behaviour of the pressure regulator valve on CNAPS is examined in Chapter
3. Analytical results are compared with data collected during tests to further refine the
models.
9
Chapter 1
XPOD Ground Support
Equipment
1. Introduction
The XPOD as shown in Figure 5 is a custom made nanosatellite separation system
designed and built at the Space Flight Laboratory. Its purpose is to secure the satellite
during the extreme conditions of the launch environment. In addition, it serves as the
interface between the satellite and the launch vehicle, and deploys the satellite upon
reaching the target orbit.
Figure 5 XPOD GNB with satellite
10
The XPOD design is derived from the previous generation of nanosatellite separation
systems called T-PODs [7]. It undergoes rigorous testing to qualify it for the space and
launch environments. However, the dynamics of deployment have not yet been
investigated in a detailed manner.
Previous deployment testing with an earlier version of the XPOD was carried out on a
bench with rollers supporting a satellite mass dummy as it exited the XPOD (Figure 6). A
mass dummy was used due to the fact that the satellite outer surface came in direct
contact with the rollers. Moreover, this kind of test only gave a rough idea about the
deployment behaviour of the XPOD; it was only possible to confirm whether the satellite
was successfully pushed out of the XPOD without measuring the deployment speed.
Figure 6 Previous setup for testing XPOD deployment
It was desired that the deployment process of the XPOD be characterized. Knowledge of
the behaviour of the satellite exiting the XPOD is useful for logistical and planning
purposes. If the path the spacecraft takes in the post-deployment period can be estimated,
this would help launch providers with placing the separation systems at appropriate
locations so as to minimize the potential for collision or interference between satellites
being deployed.
11
Given this task of characterizing the deployment of the satellite from the XPOD, a set of
requirements were defined, followed by the design and development of the ground
support equipment needed to conduct this test. A test plan was formulated and results of
the test were documented and summarized. All of this is presented below under the
relevant subsections.
2. Literature Review
Emulating zero-g conditions for the purpose of testing satellite systems has been an area
of research ever since satellite technology became more common [8]. The concept of
simulating on-orbit conditions for testing flight hardware in an Earthbound (1-g)
environment has been examined by various institutions. Due to the large amounts of
resources that are invested in designing and developing satellites, ground testing is seen
as an important step in their lifecycle. Moreover, demonstrating operations in a 1-g
environment is a positive sign that the mechanism will most likely work in the 0-g
conditions of space.
Over the years, various methods have been employed to reproduce the low-friction,
reduced gravity environment of space on Earth. These range from underwater test tanks,
free-fall tests, air bearings, and magnetic suspension [9]. During the author’s literature
review, the most common approach taken to negate friction effects was seen to be with
the use of air bearings. Lawrence Livemore National Laboratory has been using air
bearings for nearly a decade for testing the docking procedures and proximity operations
of their satellites [10]. The Lightband separation system designed by Planetary Systems
Corporation also undergoes ground testing with air bearings as part of the test setup [11].
Another method by which zero-g conditions have been emulated in a laboratory
environment, includes a robotic arm with multiple joints to provide freedom of motion in
six degrees [9], [12]. This robotic manipulator uses a sensor at the satellite-arm interface
that provides feedback to an algorithm controlling the arm.
12
3. Requirements
In what follows:
‘Shall’ implies a strict requirement
‘Should’ implies a desire
No. Requirements Description
Functional Requirements
The GSE shall provide a safe and reliable means of testing the XPOD
deployment process with a fully integrated satellite in an Earth gravity
environment. 1
Source: Reducing risk of damage to a satellite is a strict requirement.
The GSE shall be designed such that it minimizes its impact on the XPOD
deployment process. 2
Source: Large friction effects will tend to overshadow the data collected of
the deployment process.
The amount of energy introduced or lost due to the GSE on the deployment
process must be known.
3 Source: Knowledge of the influence of the GSE on the satellite during
deployment allows it to be factored out when determining the dynamics of
the deployment process on orbit.
The GSE design shall negate the effects of friction due to gravitational forces
on the deployment process. 4
Source: External forces and their effects in directions perpendicular to
motion are minimized.
13
The GSE shall allow the satellite to move freely in at least two orthogonal
directions in the deployment plane. This corresponds to motion in the X and
Y axes in the reference frame shown in Figure 8. 5
Source: Allows for studying the deployment process in a plane.
The GSE shall accommodate all predeployed spacecraft appendages.
6 Source: It is desired to test a fully assembled satellite integrated with the
XPOD.
Integration of the satellite with the GSE should require no more than 15
minutes. 7
Source: This stems from the time requirement for arming the XPOD.
Design of the GSE should be modular such that the deployment process can
be tested in at least two orthogonal planes containing the deployment axis,
which is the Y-axis as seen in Figure 8. With respect to the reference frame
shown in Figure 8, this corresponds to the XY and YZ planes. 8
Source: Future tests of the deployment process in a perpendicular plane may
be performed.
The GSE should be capable of measuring rotation rates of the satellite as it
leaves the XPOD. 9
Source: To ensure the spacecraft’s extended antennas do not collide with the
XPOD structure soon after deployment.
The height of the GSE structure from the surface on which the deployment
test is being performed shall be adjustable to +/- 5mm. 10
Source: Depending on the surface where tests are being performed, the
height may be adjusted to accommodate all satellite appendages.
14
The satellite should be supported by the GSE when it is integrated with the
XPOD. 11
Source: This ensures that the satellite is not resting on the XPOD rails, and
therefore removes the effect of friction due to gravitational effects.
12 The effect of the GSE on the satellite’s deployment speed shall be limited to
20% of the predicted on-orbit value.
Source: To gain confidence in the results of the deployment tests, the effect
of the test equipment should be limited. The 20% margin comes from the
uncertainty of the deployment speed by itself (without the effect of the GSE)
due to friction effects between the satellite and XPOD rails.
Operational Requirements
The GSE shall not interfere with the operation of the XPOD release
mechanism. 13
Source: If the release mechanism is not representative of that on orbit, the
results of the tests will be invalid.
14 The GSE carrying the satellite shall be brought to rest in a controlled manner
without risking damage to the satellite.
Source: Pertains to reducing the risk of damaging the satellite.
Structural Requirements
The GSE shall be capable of carrying a satellite weighing up to 7.5 kg. 15
Source: Stems from the XPOD GNB requirement.
16
The GSE shall hold the XPOD in a horizontal position (XY or YZ plane in
Figure 8 being parallel to surface) during tests such that the deployment axis
(Y-axis in Figure 8) is parallel to the surface and the XPOD door opens
parallel to the ground.
15
Source: To prevent the XPOD rails from coming in contact with the satellite
and imparting energy to the system. The XPOD door opening sideways helps
to ensure that there is little effect of gravity on the door opening speed.
The GSE frame supporting the XPOD shall be capable of bearing the weight
of the XPOD without deflecting more than 0.55 mm under load. 17
Source: Clearance between satellite and XPOD rails is 0.55 mm.
Future Enhancements
The GSE should have a modular design to test the CanX-4/-5 satellites being
deployed from the XPOD DUO.
18 Source: The deployment of the joined CanX-4/-5 satellites needs to be tested
to characterize the dynamics of the deployment process. Rather than
redesigning the GSE, the design should be modular so that it can be modified
to be used with the XPOD DUO and CanX-4/-5 satellites.
4. Design
The XPOD GSE for the purpose of conducting deployment tests, consist of two parts
listed below. Further details of each part are mentioned in following sections.
• The Ground Test Vehicle (GTV) is responsible for supporting the satellite during the
test. It moves along with the satellite as it is deployed from the XPOD.
• The XPOD Mount is a frame that holds the XPOD in a horizontal orientation during
the test.
16
4.1. Ground Test Vehicle (GTV)
Some of the design drivers for the design of the GTV were those concerning the reduced
friction conditions, and its compatibility with a fully assembled satellite inside the
XPOD. The method chosen to provide the reduced friction conditions are detailed in
another section below. Regardless of the method used to provide the reduced friction
environment, the test setup would require some form of supporting structure to carry the
satellite after it exits the XPOD. The main structural components of the GTV can be
categorized as the supporting structure and the parts interfacing with the satellite.
4.1.1. Supporting Structure
The primary objective of the supporting structure is to provide a platform to carry the
satellite. Some of the key requirements that governed the design of the platform were the
ability to carry a satellite weighing up to 7.5 kg, and having the height adjustable. Also,
wherever possible, commercially available parts and brackets were used in the design.
Regardless of the method used to provide the reduced friction conditions required, the
structure would include a flat plate to which the remaining parts could be fixed.
After undergoing many design iterations, the final design consisted of a number of small
parts that could be assembled using a few screws and simple tools. In order to reduce
machining costs, some of the complex designed parts were simplified into smaller parts.
Some features of the design such as the placement of screw holes were based on the
dimensions of commercially available angle brackets.
The supporting frame consists of a base plate atop which are fixed four angle-brackets.
The base plate was designed with the idea of using ball transfers as a means of reducing
friction between the GTV and surface. Justification for the use of ball transfers is
presented in a later section. Cutouts were added to reduce the mass of the plate. Each of
two C-brackets sits on two of the angle-brackets. A secondary plate is fixed to the top of
the C-brackets.
17
Figure 7 Supporting structure of GTV
At this point, it is worth mentioning which satellite panel faces the GTV. With the
satellite inside the XPOD, only the +/-Z panels are exposed as shown in Figure 8. The +Z
panel has the magnetometer boom attached to -Y end of it, while the –Z panel has the
VHF antenna on the same end. Given that the VHF antenna is not a part of all GNB
satellites, it was decided to have the satellite’s –Z panel face the GTV. Since it was
required of the GTV to be able to accommodate all fixed appendages, a cut-out was
included in the secondary plate for the antenna to pass through
Figure 8 Exposed faces of satellite when integrated with XPOD
18
A base block is the next component in the order of arrangement. But rather than resting
directly on the secondary plate, the block is held off the plate using shims and spacers as
seen in Figure 9. The number of shims can be varied to adjust the height of the satellite so
that the trays line up with the XPOD rails. Depending on the surface on which the test is
being conducted, the height of the satellite from the surface may need to be adjusted. The
centre of the base block has a hole (Figure 9), through which a screw-spacer pair passes
through and secures the satellite interface adapter (called the H-frame) to it. More details
about the H-frame and its purpose can be found in the following section.
Figure 9 Shims used to adjust height of base block
4.1.2. Satellite Interface Adapter
The key constraint on the design of the adapter that fixes the satellite to the GTV was that
it should not interfere with the XPOD’s release mechanism and the deployment process.
When the satellite has been integrated with the XPOD and the release mechanism armed,
only the +/-Z faces of the satellite are exposed. The X panels are covered by the XPOD’s
side panels, while the Y panels face the XPOD’s pusher plate and door respectively. As
mentioned in an earlier section, the –Z panel was seen as the favourable option for
interfacing with the satellite due to the fact that the VHF antenna was the only fixed
appendage on this side, as opposed to the larger magnetometer boom on the +Z side.
19
However, the absence of pick-up points on this panel prompted the conception of an
alternative approach.
It was decided to exploit the clearance between the satellite’s Y panels and the XPOD
pusher plate and door as shown in Figure 8. Although the gap existing between the
XPOD and satellite is small, having the adapter fastened to the Y faces allowed the
satellite to be picked up by its tray rather than a panel. The GNB trays serve as the
primary structure for the spacecraft and thus provide more reliable pick-up points. The
design of the adapter was thus driven by its ability to fit into the clearance that existed
between the XPOD and the satellite’s Y panels.
The ‘H-frame’, as the adapter is called, serves as the interface between the satellite and
the remaining GTV structure. It consists of an H-shaped bracket, to the ends of which L-
brackets are fastened. The H-bracket was sized so that it did not deflect more than 0.55
mm under the weight of a 7.5 kg satellite. This constraint was driven by the requirement
that the satellite be supported by the GTV when integrated with the XPOD. A deflection
greater than 0.55 mm would cause the spacecraft to come into contact with the XPOD
rails.
A hole in the centre of the H-bracket was included to meet an initial requirement for the
satellite to be able to rotate freely while on the GTV. This would have been achieved by
having the H-frame pivot about a spacer through the centre hole. The spacer in turn
would be fastened with a screw coming from under the base block as shown in Figure 10.
A washer between the H-frame and base block would allow the frame to rotate freely. For
the same reason, the width of the H-bracket’s cross-arm was made small so as to reduce
contact area between it and the block. In this configuration, the satellite, with the H-frame
fastened to it, could be lifted off the GTV without the need for any unscrewing.
20
Figure 10 Earlier design of H-frame, allowing satellite to rotate independent of GTV
The added benefit of this approach was the fact that an end-to-end test of the XPOD
deployment could be carried out; i.e. the test could be done after the XPOD containing
the satellite had been through a series of thermal and vibration tests. Analysing the
ejection process before and after the thermal and vibration tests would have provided
insight into the effects of the tests on the performance of the release mechanism. The only
drawback about this method was having the L-brackets remain attached to the satellite
during thermal and vibration tests. With the satellite inside the XPOD, only the H-bracket
would be accessible and could be taken out by unscrewing it from the L-brackets as seen
in Figure 12. Although small in mass and dimensions with respect to the spacecraft, the
presence of the L-brackets would have affected the test results, especially for the
vibration tests. Hence it was decided at the managerial level to do the deployment tests
separate from the thermal and vibration tests.
The current design has the H-frame securely fastened to the base block. The only
difference to the original assembly is that rather than having a screw coming from under
the base block, it passes through the spacer at the centre hole of the H-frame, and screws
into the top of the base block as seen in Figure 11. The clearance between the satellite’s
panel and the H-bracket is sufficient for the screw head to be protruding from the hole.
21
Figure 11 Method of fastening H-frame to base block in earlier design
The L-brackets are positioned such that they fasten on to the Y panels of the spacecraft,
facing the XPOD pusher plate and door respectively. The L-brackets were sized to fit into
the clearance between the satellite and XPOD, without coming into contact with either
one. As well, the holes on the L-brackets were positioned such that the satellite is held off
the H-frame as seen in Figure 13. Spacers between the satellite panel and L-bracket
prevent any form of contact between the two due to vibrations experienced during the
deployment process.
Figure 12 Components of H-frame
22
Figure 13 Method of attaching H-frame to satellite
When the satellite attached to the GTV is integrated with the XPOD, the L-brackets are
present in the space existing between the +/-Y faces of the satellite and the XPOD door
and pusher plate respectively.
With the exception of the angle brackets on the GTV, all parts were custom made.
Engineering drawings were drafted for each component and sent to a local machine shop
for fabrication. Aluminum alloy 6061-T6 was the material chosen for all custom-made
parts of the GTV due to its ease of availability as well as reduced machining costs. The
parts were appropriately sized to meet the deflection requirements.
4.2. Reduced friction condition
Since one of the driving requirements was to provide reduced friction conditions for the
satellite as it leaves the XPOD, different methods of meeting this requirement were
considered. This feature would be part of the GTV base plate.
23
• One option was to make use of a material of which the surface properties were known
and well understood. Delrin was seen to be a good choice since it is used inside the
XPOD, where the satellite rails interface with it. The friction coefficient between
aluminum and Delrin was determined in previous tests during the design of the
XPOD. Knowing the friction between the aluminum base plate and Delrin sheet, the
acceleration of the GTV and satellite could be determined, from which the satellite’s
exit speed from the XPOD could be calculated.
• On the other end of the spectrum of providing reduced friction conditions for the
satellite and GTV, the concept of air cushioning was considered. Either by having an
air table forcing out compressed air from its surface through little holes, or by having
compressed air forced out form the bottom of the GTV similar to a hovercraft, the
contact between the table surface and the base plate would be minimized, thereby
making the effects of friction negligible. The GTV would be free to move about in a
plane, and free to rotate about an axis perpendicular to that plane.
• Considering more commercially accessible and off-the-shelf hardware, low friction
bearings were seen as another alternative to provide reduced friction conditions for
the deployment tests. Stainless steel low friction ball transfers were purchased and
tested in the lab to determine the friction effects between the transfers and a smooth
surface such as a tiled floor. The loads on the ball transfers were representative of the
satellite and GTV.
The criteria used to evaluate each of the above three options for providing low friction
conditions stem from some of the requirements mentioned in an earlier section. They are
listed in Table 1 where the following options are defined:
Option A: Sheet of low friction material for GTV to move on
Option B: Low friction ball transfers fastened to GTV
Option C: Air cushioning
24
Table 1 List of criteria used to evaluate options for reducing friction effects
WeightCriteria Rating Score Rating Score Rating Score
1 Reliability/Safety 3 5 15 5 15 1 32 Implementation 1 5 5 5 5 3 33 Friction reduction 3 1 3 3 9 5 154 Ease of repeatability 2 5 10 5 10 1 25 Range of Motion 2 5 10 5 10 3 6
Total 43 49 29
Option A Option B Option C
Weighting scale:
1=Not very important
2=Important
3=Very important
Rating:
1=Low
3=Medium
5=High
4.2.1. Criteria Weighting and Option Rating
1. Reliability/Safety
Safety of the satellite during tests is a key requirement of the XPOD GSE. Since tests will
include a fully assembled satellite, it is important to reduce all potential risks of
damaging the spacecraft to a negligible level.
Option Rating-Options A and B offer safe methods for allowing the GTV to move as it
exits the XPOD. Option C was given a rating of 1 because of the mechanism by which air
bearings work which does not make them very reliable on a relative scale. Air bearings
force the GTV off the surface so that it is suspended on a cushion of air. Therefore they
require a continuous and uninterrupted supply of pressurized air for extended periods of
time. If the air supply gets obstructed or is exhausted, this might cause the GTV to ‘drop’
onto the surface.
25
2. Implementation
This criterion refers to implementing the suggested option for reducing the effects of
friction. It does not influence the outcome of the test but merely quantifies the amount of
effort and resources to be utilized to employ a particular method. Thus, it was only given
a weight of 1.
Option Rating-Options A and B are relatively easier and simpler to implement. Option A
requires laying out a flat sheet of Delrin on which the GTV will move. For option B, ball
transfers have to be fastened to the bottom of the GTV. Air bearings on the other hand
require more extensive preparation in terms of arranging for means to supply compressed
air to the bearings at a constant flow rate, which is why it was given a lower rating.
3. Friction Reduction
Again, this is one of the requirements of the XPOD GSE. Reducing the effects of friction
allows for testing the deployment in conditions that are closer to the ideal case on orbit. A
large friction force will eclipse the force applied by the XPOD spring, not allowing the
dynamics of the deployment process to be captured to a sufficient level of detail.
Option Rating-Having the GTV in contact with the surface, as in option A, produces a
relatively large friction force due to the larger friction coefficient between Delrin and
metal compared to the ball transfers. In the case of option C, the GTV sits off the surface,
suspended on a cushion of air, thereby reducing the effects of friction to nearly zero.
Option B, ball transfers, falls in between the two extremes, reducing friction significantly,
but not completely eliminating it.
4. Ease of Repeatability
Since a number of tests are required to obtain an average deployment speed, it was
considered fairly important that the test be easy to repeat. If a long time was needed to set
up the equipment for the test, this directly affected the number of tests that could be done
in a given time.
26
Option Rating-Use of air bearings would require that the bearings be inspected and the
supply tank refilled on a frequent basis. On the other hand, options A and B require
minimal preparation, allowing successive tests to continue uninterrupted. Wear and tear
of the Delrin and ball transfers can be addressed by simply replacing them on a regular
basis.
5. Range of Motion (Translational)
The extent to which the GTV would be allowed to move was also considered to be of
some significance. It is required that the test capture the satellite’s translational motion
following deployment. Therefore, the GTV would need to move freely a short distance
after leaving the XPOD.
Option Rating-The Delrin sheet and ball transfers do not limit the range of motion of the
GTV in any manner. However, using air bearings may constrain the range of the GTV,
depending on the method by which air is being supplied to them: if a hose is connected to
the GTV to supply the pressurized air, the range is limited by the length of the hose; in
the case where the compressed air supply is stored on the GTV, the range is limited by
the capacity of the tank because if the air runs out, the GTV will drop to the surface and
remain stationary.
Based on the ratings and total score of the decision matrix, ball transfers were chosen as
the method of reducing the effects of friction on the GTV. Commercially available ball
transfers were obtained with threaded studs which screw directly onto the base plate of
the GTV. To determine the friction coefficient between the ball transfers and surface, a
series of tests were done where a known force was applied to the GTV over a fixed
distance. By equating the known and unknown (friction) force acting on the GTV to the
net acceleration, the friction coefficient was derived. This coefficient was used for the
purpose of creating an analytical model to simulate the deployment of the satellite from
the XPOD.
27
4.3. XPOD Mount
During testing, it is required that the XPOD be held in a horizontal orientation, with the
door opening sideways. This ensures that the entire test is carried out in 1-g conditions,
and therefore the effect of gravity on the test can be neglected. The safest and most
reliable way of securing the XPOD was seen to be through the base plate. The XPOD
base plate has been designed to fasten on to the secondary payload platform on the launch
vehicle. It consists of a hole pattern across its entire width and length for up to 32 screws.
Therefore, the design of the XPOD Mount focused on securing the XPOD through its
base plate.
The XPOD Mount is a stationary frame meant to hold the XPOD in a horizontal
orientation during deployment tests. It consists of two plates perpendicular to each other;
one lays flat against the surface, while the XPOD affixes to the vertical plate. The XPOD
base plate is screwed onto the vertical plate which has the same bolt pattern on it as the
XPOD base. Angle brackets screw into both plates and provide the reinforcement to hold
them in place.
The Mount was designed to have a large mass, first to provide a stable base for the
XPOD, and second to dampen any vibrations during testing that may arise due to the
recoil action of the spring. A moment analysis was carried out to determine the minimum
mass of the horizontal plate required to prevent the XPOD and Mount from tipping over.
The GTV is designed so that the satellite’s height with respect to the surface can be
adjusted such that the satellite rests on the GTV rather than on the XPOD rails. But for
the sake of the moment analysis, in the interest of being conservative, it was assumed that
the spacecraft is resting on the XPOD rails.
28
Figure 14 Tipping analysis for XPOD fastened to Mount
Considering moments about point A, with counter-clockwise as positive, it was
determined that the mass of the base plate is sufficient to keep the Mount or XPOD from
tipping over.
( ) ( ) ( )( ) ( )
( )( ) ( )( )kg
m
mkgmkgm
r
mrmrm
mrgmrgmrgM
plate
plate
satXPODplate
satXPODplateA
5.1519.0
21.05.714.08.9
0
min,
min,
≈+=
+=
=−−=∑
The minimum mass of the base plate that would prevent the XPOD from tipping over
was found to be approximately 15.5 kg. The actual mass of the base plate is 31 kg which
gives an acceptable factor of 2.
29
4.4. Mass Model
To assess and resolve any issues related to testing with the GTV, a mass model of the
satellite was used for initial tests. The mass model consists of two identical pieces that are
fixed together with some overlap. When put together, the external dimensions of the mass
model are exactly those of the GNB satellite. But the internal layout of the mass model
differs from the satellite. Each half is basically a ‘box’ like structure as shown in Figure
15 that has been hollowed out on the inside to get the desired mass. The ‘lips’ along the
open edges of each half allows for some overlap between the two pieces, ensuring they
are properly aligned. Each half of the mass model is made of aluminum 6061-T6,
weighing approximately 3.25 kg. The two halves are fastened using 20 screws around the
perimeter of the interface. Extra mass can be added to the mass model on the inside using
four screws for each half. This allows the total mass and inertia properties of the model to
be very coarsely altered to simulate a heavier GNB satellite. While the mass model has
other uses for testing of other systems, it was also used to characterize the GTV prior to
beginning the deployment test with a fully assembled satellite.
Figure 15 Mass model halves showing optional masses inside
30
4.5. GNB Spare Structure – Satellite Model
In addition to testing the GTV with the mass model, it was desired that a spare GNB
satellite structure be used for calibration purposes. This consists of the trays and panels
from the satellite, minus the internal and external components. Mass models of internal
components were designed such that the dimensions remained approximately the same,
allowing existing screw holes on the structure to be used for fastening them (Figure 16).
External protruding appendages, such as antennas were fixed to the outer panels. The
satellite model allows for testing the clearances between the fixed appendages and the
XPOD during the deployment process. It serves as a middle step between testing the
GTV with the mass dummy, and with the fully assembled satellite.
Figure 16 Mass models of components inside GNB tray
5. Mass Budget
Although no mass requirement was set, an effort was made to minimize the mass of the
GTV to reduce its effect on the satellite’s motion. A heavier GTV for a given deployment
31
force imparted by the XPOD spring meant that the satellite would experience less
acceleration. Therefore, the mass of each component of the GTV was recorded (Table 2),
and where possible, parts were made lighter either by using a different material or by
adding cutouts.
Another advantage of keeping the GTV mass low was realized later on during analysis. A
light weight GTV meant that the centre of gravity of the combined satellite and GTV
system would not shift too far from the centre of gravity of the satellite itself. Having the
centre of gravity move too far would cause the deployment force of the XPOD spring to
produce a moment about the common centre of gravity, thereby making the satellite
rotate while inside the XPOD, leading to the possibility of jamming.
Table 2 Mass budget of GTV
Part Material Mass/unit (g) Quantity Mass (g)Ball Transfer Stainless Steel 58 4 232Base Plate Al 6061-T6 702 1 702
Angle bracket Galvanized Steel 108 4 432C-bracket Al 6061-T6 77 2 154
Secondary plate Al 6061-T6 149 1 149Base block Al 6061-T6 86 1 86
H-frame Al 6061-T6 139 1 139L-bracket Al 6061-T6 4 4 16
Total 1910
6. Safety and Reliability of GTV
The GTV is a stable platform that supports the satellite once it exits the XPOD. This
section looks at the integration of the satellite with the GTV and what safety measures are
taken to reduce the possibility of damage to the fully assembled satellite.
� The interface between the satellite and GTV is solely through the L-brackets on the
H-frame. Four #4-40 screws secure the satellite in place. The screws fit into through
32
holes on the L-brackets, pass through holes on the +/- Y panels, and screw directly
into the +/- Z trays of the satellite. An unthreaded female spacer between the satellite
and L-bracket prevents any vibration or movement of the satellite once the screws are
in place. Each L-bracket is fastened to the H-bracket using two screws.
� Each of the remaining parts of the GTV is fastened to neighbouring parts with at least
two screws. The thickness of each of the plates has been sized so that the deflection
under load is within acceptable limits. This limit was determined to be less than 0.55
mm of total deflection since this is the clearance that exists between the satellite rails
and the XPOD. If the GTV supporting structure was to deflect 0.55 mm or more, the
satellite would come into contact with the XPOD as soon as the preload was
removed. This is undesirable since the contact would introduce friction, which in turn
would have to be accounted for when analyzing the results of the deployment tests.
� The ball transfers used on the GTV base plate are made of stainless steel. They
consist of a main ball that rotates on a layer of support balls housed in a cup. The
arrangement allows the main ball to rotate freely in all directions and reduces the
possibility of the ball transfer jamming up.
� Since the GTV is approximately a third of the weight of a fully assembled GNB
satellite, the centre of gravity of the combined satellite and GTV system is quite high
with respect to the surface on which the GTV rests. A tipping analysis was done to
determine the minimum speed that would cause the satellite to tip over if the GTV
was to hit an obstruction and come to an immediate stop.
An energy balance approach was employed for the analysis. By comparing the initial
kinetic energy of the GTV and satellite to that required to tip them over, the minimum
speed was calculated. The limiting case for the tip-over was seen to be when the
satellite’s centre of mass reaches the maximum height; i.e. all of the translational
kinetic energy of the system converts to potential energy (Figure 17).
33
Figure 17 Tipping analysis of satellite and GTV
( )( )smv
mms
mghv
ghv
EE
CM
CM
CM
potentialkinetic
/89.0
4081.922
2
1
2max
max2
≈
==
=
=
This means that for speeds of 0.89 m/s or greater, it is likely that the GTV will tip
over. As is mentioned in a later section, the maximum measured GTV speed on
exiting the XPOD was seen to be approximately 0.9 m/s. Considering the fact that the
speed rapidly decreases due to the presence of friction between the GTV and surface,
the potential risks associated with the GTV tipping over and damaging the satellite
were considered negligible.
� To reduce the possibility of any obstruction in the path of the GTV, a sheet of plastic
is used to provide a smooth even surface to roll on. The length of the sheet was
determined by modelling the dynamics of the satellite and GTV system taking into
account the deployment speed from the XPOD and friction forces acting on the ball
transfers. Before the test, procedures require that the plastic sheet be cleaned and
inspected for cracks and protrusions that would interrupt the motion of the GTV.
34
7. Data Collection Methods
Data from the deployment tests is collected by a three axis accelerometer and single axis
gyroscope shown in Figure 18, and is transmitted wirelessly via a Bluetooth connection.
A USB dongle connected to the computer receives the transmitted data in terms of the
gravitational acceleration, g. These values are then used to calculate the speed of the
satellite as it leaves the XPOD.
Figure 18 Accelerometer used for deployment tests (courtesy SparkFun Electronics)
8. Deployment Analysis
The XPOD separation system was designed to deploy the spacecraft into orbit at a certain
relative velocity. Hence, the XPOD design takes into account the different spring
constants for the door hinge and the main deployment spring, to ensure that the door is
fully open when the satellite begins to be pushed out. Therefore the analysis below
assumes that the door is completely open and not in the path of the satellite as it is being
deployed.
35
Prior to the activation of the release mechanism on the XPOD, the satellite’s motion in
the longitudinal and lateral directions is limited due to the preload that boxes it between
the pusher plate and door. When the signal is received from the launch vehicle, the
release mechanism activates, and the door holding the satellite inside, swings open. At
this instant, the preload provides a large acceleration spanning a short duration. This
initial ‘push’ is meant to set the satellite in motion and overcome any initial resistance to
the deployment due to static friction on the satellite inside the XPOD.
Following this, the main spring provides the remaining energy to push the satellite out of
the XPOD. The length of the spring is slightly greater than the length of the satellite from
rail tip to tip. This means that more than half the satellite is out of the XPOD before the
main spring force goes to zero.
In the ideal case, the satellite continues to move away from the XPOD due to the energy
imparted to it from the spring. With no external forces or torques acting on it, the
spacecraft including all fixed appendages clear the XPOD structure within a few seconds,
depending on the exit velocity.
With respect to the deployment tests performed in 1-g conditions, a few other factors
come into the picture that would otherwise have not been an issue on orbit. The fact that
there is some form of contact between the GTV and surface gives rise to friction effects
which need to be accounted for. In addition, depending on how the spacecraft sits inside
the XPOD prior to deployment, the friction forces at the satellite-XPOD interface also
have to be considered. Furthermore, because the GTV is attached to the satellite during
the test, the energy imparted by the main spring is used to move the mass of the GTV and
the satellite. These effects and their influence on the spacecraft’s motion are elaborated
on below.
The preload provides a rapid increase in acceleration to the satellite attached to the GTV.
However this force acts over a very short distance and time frame. Following this
occurrence, the main spring force becomes the dominant force, pushing more than half
36
the satellite out of the XPOD. Up until this point, the friction between the GTV and the
surface is insignificant, being an order of magnitude less than the dominant spring force.
Once the spring has fully extended to its resting length, the only force acting on the
satellite and GTV is friction. When the friction force is factored into the equations of
motion, the exit velocity of the satellite in the ideal case can be predicted.
An analytical model was created to simulate the motion of the satellite and GTV during
deployment. Predicted exit speeds are shown in the graph below in Figure 19, along with
a plot for the case without the presence of the GTV affecting the satellite’s motion.
Figure 19 Estimated deployment speeds of satellite based on analytical model
The above graph shows velocity profiles for three cases.
• The yellow line depicts the theoretical velocity profile of the satellite from its rest
position before deployment. This model assumes that the satellite remains centred
within the XPOD, and therefore has nothing opposing its motion. The satellite is
37
accelerated by the spring force, and continues moving away from the XPOD at a
constant speed.
• From a more realistic perspective, the satellite will come into contact with the XPOD
rails due to the vibration experienced when the door swings open. This causes a small
friction force to exist at the satellite-XPOD interface, acting in the opposite direction
to the spring force. The friction effects are relatively small when compared to the
force exerted by the main spring. Only when the spring force goes to zero does the
friction force become the dominant force acting on the spacecraft. It causes a gradual
deceleration of the satellite until it has completely exited the XPOD, as is shown by
the blue line in the graph. The analysis for this case was done by Stephen Mauthe
when determining the spring constant for the main spring [7].
• The theoretical and realistic cases discussed above encompass the two extreme cases
that the satellite may experience during a successful deployment from the XPOD.
However, in the case of ground testing of the deployment process, the influence of the
GTV on the spacecraft’s speed had to be accounted for. Using the empirical friction
coefficient values for the ball transfers, the satellite’s speed was predicted, as
represented by the pink line in the above graph. With the satellite resting on the GTV,
the friction effects at the satellite-XPOD interface are removed since the spacecraft no
longer rests on the XPOD’s rails. The friction between the ball transfers and surface
was determined to be a magnitude less than that between the satellite and XPOD.
This is evident in the plot, which shows that after the spring force is removed, the
GTV and satellite decelerate at a rate slower than for the realistic case (blue line)
without the GTV.
38
9. Testing
Deployment tests were carried out with an XPOD that had undergone vibration testing.
All test equipment was setup on the floor on top of a plastic sheet as shown in Figure 20
which provided the smooth surface for the GTV to move on. The mass dummy was used
for the initial tests. XPOD arming procedures, as outlined in [14], were followed to arm
the XPOD’s release mechanism for each test. This was to ensure that the appropriate
preload was applied on the mass dummy to represent flight conditions. Test procedures
outlined in [15] were followed.
Figure 20 Setup of GSE for deployment tests
Data collected with the accelerometer was found to be noisy due to the vibration
experienced when the XPOD door swings open and also because of the small vibrations
from the ball transfers when the GTV is in motion. To overcome this, a moving average
of the collected data was used to filter out noise, resulting in a ‘cleaner’ profile of the
speed (Figure 21).
39
Figure 21 Sample plot of data collected by accelerometer during test
XPOD Deployment Speeds
0.7
0.75
0.8
0.85
0.9
0.95
1 2 3 4 5 6 7 8 9 10 11 12 13 14 15
Test #
Spe
ed (
m/s
)
Figure 22 Deployment speeds measured during tests
The deployment speed of the satellite measured over a series of tests was found to be
within 10% of the theoretical value of 0.94 m/s. Inconsistencies in measured speeds of
successive tests was mainly due to the arming procedure of the XPOD; despite the best of
efforts, no two arming processes were the same. No rotation of the satellite and GTV was
40
observed or measured. Lateral motion in the X-axis (Figure 8) was negligible; the
satellite was seen to move less than 5 cm in the X-axis after exiting the XPOD. Some of
the reasons for the discrepancies between the predicted speed and those measured are as
follows:
• Uneven surface: During testing, it was noticed that the floor was sloped
approximately 1 degree in the direction of deployment. Although this slope is not
sufficient for a component of the gravitational force to accelerate or decelerate the
GTV, it causes the XPOD Mount to tilt, which prevents the XPOD from remaining
horizontal to the floor. This in turn leads to the satellite coming into contact with the
XPOD rails, increasing the resistive force opposing forward motion.
• Preload: To reduce risk to an XPOD meant for flight, the engineering model XPOD
was used for the purpose of the deployment tests. This XPOD had already been
through vibration and thermal testing which caused some misalignment in the
structure, most noticeable in the door. When the XPOD was armed, the door did not
close completely as a result of the misalignment, which, because of the XPOD design,
reduced the preload acting on the spacecraft. Therefore, the initial ‘push’ that the
satellite experiences when the door opens may have been less than what it was
supposed to be.
• Arming Procedure: The only variable for each test was the manner in which the
XPOD was armed. Although care was taken to follow the procedures exactly as
outlined, the release mechanism’s vectran cord may not have been pre-tensioned to
the same level each time due to the author’s limited experience in arming procedures.
This in turn would have indirectly affected the pre-load.
• Friction Coefficient: Perhaps one of the biggest uncertainties is the friction coefficient
that was empirically derived for the ball transfers, and used to predict the influence of
the GTV on the satellite’s deployment.
41
10. Conclusion
Despite the discrepancies, the measured speeds were found to be within limits of the
requirements. The average speed of 15 full-scale deployment tests was approximately
0.85 m/s, giving an error within 10% of the predicted value. This shows that the
dynamics of the satellite being deployed from the XPOD are well-captured by the
analytical model, and can further be extrapolated to estimate the satellite’s path.
Future tests can analyze the dynamics of deployment in the YZ plane (refer to Figure 8),
as well as rotation about other axes to better characterize the deployment process. Results
can be compared to the on-orbit case to determine how well the GSE is capable of
capturing the ideal case deployment dynamics.
Moreover, in the long term, some steps could be taken to reduce the influence of external
sources on the measurements. These include using a fully functional XPOD that hasn’t
been put through vibration testing, significantly reducing friction effects by using air
bearings, using a higher resolution accelerometer.
42
Chapter 2
Thrust Stand Design and
Calibration
1. Introduction
The CanX-4/-5 mission is meant to demonstrate formation flying capability of
nanosatellites in orbit. The two satellites, designated chief and deputy, use a series of
discrete thrust schemes to enter into and maintain formation flight. Thrust is produced by
the onboard propulsion system, Canadian Nanosatellite Advanced Propulsion System
(CNAPS). An algorithm running on one of the onboard computers, computes the required
thrusts for a desired manoeuvre to maintain formation.
As robust and parameterized as the formation flying algorithm is, the ultimate goal of
demonstrating precision formation flying relies largely on the performance of CNAPS. In
other words, the question that arises is whether CNAPS will perform as designed and
expected. Although it is a derivative of the previous generation of SFL’s propulsion
system, NANOPS, the requirements for CNAPS are far more stringent and its
performance of greater significance to the mission.
Rather than having a single nozzle like NANOPS, CNAPS employs a four nozzle system
primarily to reduce momentum build-up in the satellite due to pointing errors. While this
also provides redundancy, it also requires further testing. The performance of each of the
43
four nozzles needs to be characterized so that any inconsistency between measurements
and predictions are better understood.
This section describes the design of a thrust stand for the purpose of characterizing the
performance CNAPS. As well, calibration of the stand and test procedures are outlined.
2. Literature Review
Recent advances in technology have afforded the implementation of propulsion systems
on micro- and nano-scale satellites. These systems are capable of producing thrusts on the
order of milli-Newtons, and therefore require different test setups and equipment for
calibration and ground testing. Knowledge and control over the thrust profile is of
significant importance when dealing with relatively small satellites to prevent undesired
momentum build-up and attitude changes.
Most existing thrust stands employ the use of moments induced by the thrust force to
measure the performance of the propulsion system. The types of thrust stands found
during a literature survey may be categorized into the vertical and horizontal forms.
Vertical thrust stands are pendulum-like in appearance, having a rod pivoted on one end,
while thrust forces are applied, and measurements made at the other end [16]. Stands of
the horizontal form include a beam balanced on a fulcrum. Thrust forces and
measurement equipment are placed at opposite ends of the beam.
Typically, measurement of the thrust force is made by noting the displacement of the
thrust stand from its equilibrium position. Linear variable differential transformers
(LVDT) have been used for this purpose [17]. Electrostatic combs have also been
employed in cases where higher resolution of the thrust profile and impulses is desired
[18].
44
3. Requirements
In what follows:
‘Shall’ implies a strict requirement
‘Should’ implies a desire
No. Requirements Description
The thrust stand shall be capable of measuring a minimum and maximum force
of 5 mN and 200 mN respectively. 1
Source: Depending on the secondary tank pressure and the number of nozzles
in use, the thrust magnitude ranges from 5 mN up to 200 mN.
The stand shall be capable of measuring the thrust force to a resolution of at
least 0.5 mN and should be able to measure with a resolution of 0.1 mN. 2
Source: The formation flying algorithm imposes a tolerance on the precise
knowledge of the thrust magnitude.
The stand with CNAPS attached to it shall be able to fit and operate inside
SFL’s larger vacuum chamber, MIR. 3
Source: Test plan for CNAPS includes testing in vacuum conditions [19].
The stand shall have the ability to fasten CNAPS to it in a secure manner.
4 Source: CNAPS should be rigidly fastened to the thrust stand to accurately
detect thrust forces and to reduce risk of damage.
The stand shall be capable of measuring thrust and impulse at a frequency of at
least 5000 Hz. 5
Source: The solenoid valves used have a maximum frequency of 500 Hz [20].
45
Standard practice is to measure at 10x component frequency.
The thrust stand shall measure the amount of fuel used during tests.
6 Source: Determination of specific impulse requires knowledge of mass of fuel
consumed.
The thrust stand shall measure the thrust misalignment up to 5 degrees.
7 Source: The formation flying algorithm imposes a tolerance on the knowledge
of the thrust direction.
4. Design
The requirement for a thrust stand derives mainly from the more stringent performance
requirements for CNAPS with respect to those for NANOPS. Ground tests for CNAPS
require it to be tested to a higher level of accuracy and resolution. The previous method
used for testing NANOPS was deemed insufficient to measure the mN thrusts to a µN
resolution and short duration impulses produced by CNAPS.
A literature review suggested that the most common method for measuring mN forces
was with test stands that amplified the small forces produced [16], [17], [18]. Beginning
with this idea, the next step was to determine a way to measure the force produced.
Previous thrust stand designs made use of electro-static combs (ESC). Although this
method gave a high degree of accuracy for even nN forces, the effort required in
calibrating the ESC was considered to make it inefficient for the given task.
The next option was to research force sensors capable of measuring forces to a resolution
of µN. Much of this surveying was done by a previous SFL student, Jonathan Grzymisch
and SFL staff member, Stephen Mauthe and the author is appreciative of their efforts.
46
Eventually, the choice of using a sensor for measurement, and selection of the particular
sensor was made at a managerial level. This had a small impact on the constraints
imposed on the thrust stand design, mainly concerning the measuring range and limits of
the sensor.
The force sensor acquired is a COTS component, in keeping with SFL philosophy. Two
of the key features of the sensor that had to be kept in mind when designing the thrust
stand, were the full scale measuring range and the resolution. The measuring range of the
sensor was important because this set the upper limit of the amount by which the thrust
forces could be amplified. The sensor resolution, which is its ability to detect the smallest
force, dictated the lower limit of the force amplification.
The remaining design of the structure of the thrust stand was governed by CNAPS
requirements and limitations of the sensor. It was decided to make use of moment arms as
a means of amplifying the small thrust forces to a level that could be measured by the
force sensor. This would be achieved by having CNAPS and the sensor fastened to two
ends of a pivoted beam. In sizing the dimensions of the beam, conflicting requirements of
the sensor’s measuring range and the ability to detect forces to a resolution of µN had to
be kept in mind. While some thrust stands were found to have a vertical design during the
literature search, the added requirement of having the stand fit inside SFL’s vacuum
chamber forced the design to be horizontally oriented. A derivative of this requirement
was the fact that because a single door to the vacuum chamber provided limited access
and reach inside, the maximum length of the beam would be restricted to one’s ability to
reach it when it was fully setup inside the chamber.
The thrust stand consists of a ‘multiplier arm’ as the beam is called, pivoted about a male
pivot block as shown in Figure 23. The pivot design was chosen to minimize the amount
of friction seen by the multiplier arm during rotation. A female pivot block on the
underside of the arm provides the interface between the arm and the pivot. The length of
the pivot blocks was sized to prevent rotation about a vertical axis (Z-axis of thrust stand
reference frame, Figure 31) through the pivot. The male portion of the pivot is fixed onto
47
a support block, which rests on a base plate. The base plate was sized so that it fits
breadth-wise inside surface of the vacuum chamber. Delrin blocks on the underside of the
plate keep it off the vacuum chamber surface, thereby providing some form of insulation
from background vibrations.
The half-angles of the male and female pivot blocks are 45 degrees and 55 degrees
respectively. Therefore, the only contact that remains between the two blocks is at the
vertex.
Figure 23 Fulcrum for thrust stand provided by pivot blocks
In designing the multiplier arm, the length was determined by calculating the
amplification of force required such that it could be measured by the force sensor. Since
the sensor can only measure forces to a resolution of 1 mN, the distance from the point
where CNAPS is fastened, to the pivot would need to be at least 10 times greater than the
distance between the sensor and pivot. This ratio of moment arms would amplify the sub-
mN level thrust force from CNAPS to the mN level.
While the above provided the lower limit for the length of the multiplier arm, the upper
limit for its length was constrained by two things: the maximum force that the sensor
could sustain without incurring permanent damage; and the ability to reach the end of the
arm furthest from the door to the vacuum chamber while the test setup was inside it.
Given that CNAPS fully fuelled weighs approximately 17 N, which when multiplied with
a moment arm ratio of 10 results in a force much higher than the sensor’s limit, it was
48
evident that a counterweight would be needed. The counterweight would serve to remove
the effect of the weight of CNAPS. Moreover, because it was initially believed that the
sensor could only measure in compression, the counterweight was appropriately sized so
that the sensor remained in compression even with CNAPS’ fuel tanks emptied. Figure
24 and Table 3 show the moment arms of the applied forces with respect to the pivot.
Figure 24 Schematic of thrust stand identifying forces and their locations
Table 3 Sizing of multiplier arm based on force sensor limits
Counter weight
Arm
Max Min Max MinMass (g) 5000 255 1775 1475Force (N) 49.1 2.5 17.4 14.5 0.16 0.005
Moment arm wrt pivot (mm)
91.5 84 305 305 305 305
Moment (N.mm) 4488 210 5311 4413 49 2
Max force measured by sensor 43.3 NSensor Upper Limit 50.0 NMargin 13.5%
Min force measured by sensor 5.5 NSensor Lower Limit -50.0 N
CNAPS Thrust
49
The multiplier arm was sized based upon Table 3. The moment arm of each applied force
was varied so that the resulting force on the sensor was within its limits. The moment
equation used is as follows, with counter clockwise moments assumed to be positive:
( ) ( ) ( ) ( ) ( )∑ +−−+= ThrustCNAPSarmsensorghtcounterweipivot FrFrFrFrFrM
One end of the arm has a slight protrusion; this is the point for fastening CNAPS to the
thrust stand. The other end of the arm has two holes, through which two threaded steel
rods pass through. These rods act as rails for the counterweights to slide through. The
arm is pivoted to one side of its centre to create the appropriate moment arm ratio for
CNAPS and the sensor, with respect to the pivot. To better amplify the mN thrust
produced by CNAPS, the moment arm ratio for CNAPS and the force sensor was chosen
as ~12. This was the maximum allowable ratio so that the requirement that both ends of
the arm be accessible inside the vacuum chamber was satisfied. The only difference this
would make is to the amount of counterweight that would be needed to balance out the
arm.
Figure 25 Features of multiplier arm
An adapter bracket is attached onto one of CNAPS’ side frames, which serves as an
interface between the propulsion system and the remaining thrust stand. The adapter is
then fastened to the end of the multiplier arm with screws. The adapter is a rectangular
frame with holes that line up with those on CNAPS’ side frame. A protrusion in the
centre of the adapter provides some clearance between the underside of the arm and
50
CNAPS. On CNAPS, the side opposite to the one with the nozzles was chosen to attach
the adapter so that any interference with the nozzles or the flow escaping would be
avoided. In this configuration, CNAPS is positioned with its X-axis pointed vertically up
as shown in Figure 26.
Figure 26 Method of fastening CNAPS to thrust stand
Two additional adapter brackets were designed and built. These are used to fasten
CNAPS to the thrust stand in perpendicular orientations to measure the thrust
misalignment angle. The calculations concerning the determination of the misalignment
are covered in a later section. These brackets orient CNAPS such that its Y-axis and Z-
axis respectively are pointed vertically up or down, as seen in Figure 28. Adapters for
both Y and Z-axis orientations use the same H-frame but different L-brackets, the
difference being the spacing between the holes used to fasten CNAPS in each orientation
(Figure 27).
51
Figure 27 H-frame used for measuring misalignment of thrust vector
Figure 28 CNAPS configuration for measuring thrust vector misalignment
Counterweights were custom made in denominations of 1 kg and 100 g in the form of
rectangular blocks and plates respectively. Stainless steel was chosen as the material for
the weights due to its high density. Each block and plate has two holes on its largest side
as seen in Figure 29 which help align the weights with the rods at the end of the arm, and
52
prevent any movement. This helps to ensure that the moment arm between the
counterweights and the pivot remains constant.
Figure 29 Adding counterweights to multiplier arm
Stop blocks placed under the multiplier arm as shown in Figure 30 limit the rotation of
the arm so as to avoid damage to the sensor. The height of the blocks was carefully
selected by taking into account the extension of the force sensor due compressive and
tensile loading. They are fastened to the base plate and sit approximately 0.2 mm under
the bottom of the arm when it is level. Calculations and analysis showed that for the
measuring range of interest, the total deflection of the arm and sensor will be less than 0.2
mm. The stop blocks are useful during setting up the thrust stand and provide a reference
frame to check if the arm has been balanced properly.
53
Figure 30 Thrust stand with CNAPS attached to it
5. Data Collection
Apart from the force sensor, there are other components used to collect data from the
thrust stand. The force senor, being a transducer, outputs a charge. An accompanying
charge amplifier converts this charge to a voltage which can then be read by any voltage
measuring equipment. Initial attempts with an oscilloscope and hand-held multimetre did
not give the desired temporal resolution.
Since the thrust stand is required to measure thrust at a frequency of 1 kHz, a brief search
on data collection equipment, concluded in the decision to acquire a DAQ. The features
of the DAQ were based on the measuring range and resolution required of the output
voltage from the charge amplifier. The calculations showing the required number of bits
for the DAQ are shown below.
To determine the maximum measuring range, the maximum change in force, as measured
by the sensor, was first computed. Assuming a mass flow rate from CNAPS of 0.1 g/s for
54
70 seconds1 and with all four nozzles active, each producing a maximum thrust of 20
mN, this translates to a force of approximately 1.8 N as measured by the sensor due to the
amplification by the multiplier arm. The charge amplifier can be reset for each test,
therefore zeroing the output, to prevent saturation of the amplifier.
( )
( )mN
mm
mmmNmNF
r
rFF
mNmNF
mNs
ms
g
sensor
sensor
CNAPSThrustsensor
Thrust
181825
3058069
F
80420
6981.970g1.0F
timemF : test70s singlefor propellant of in weight Change
max,propellant
max,
2propellant
propellant
≈×+=∆
×+∆=∆
=×
≈××=∆
××=∆ ɺ
Using the lowest possible scaling factor on the charge amplifier of 0.2 N/V, this gives a
measuring range of 9000 mV.
VVN
mN1.9
2.01818 ≈
The resolution of the sensor in voltage for the given scaling factor is:
mVVN
mN5
2.0
1 =
Therefore, for the range and resolution shown above, it was determined that a DAQ with
at least 11 bits would be required.
11
25
1.9
≈
=
bitsmV
V bits
Due to its cost effectiveness, and because it could potentially be used for other tests, a 16
bit DAQ was acquired by SFL. Connections between the sensor, charge amplifier, and
DAQ were made with high insulation co-axial cables.
1 The reconfiguration manoeuvre for CanX-4/-5 is the only time during the mission when CNAPS is fired for a period of up to 70 seconds (from discussions with Jesse Eyer and Niels Roth)
55
6. Analysis & Calculations
The following section describes the calculations associated with using the thrust stand to
measure thrust forces from CNAPS. The reference frame in black is that of the thrust
stand, fixed at the pivot point. The CNAPS reference frame is shown in blue. It is
assumed that the design of the pivot block prevents any force from being applied on the
sensor in the Y-axis (thrust stand reference frame). The general equation for the sum of
the moments about the pivot in each case is derived as follows.
( )TSTSCNTSTS
TSTSCNTSTS
STTSS
TTSTCNTS
TTSC
TTSO
TTS
STTSS
TTST
TCNC
TTSO
TTS
SSTC
FrFCrM
FrFrM
FrFrM
ℑ×ℑ+ℑ×ℑ=ℑ
ℑ×ℑ+ℑ×ℑ=ℑ
=×+×=∑
,
0
Eq. 1: ( )TSTSCNTSTS SSTCNTSCO FrFCrM ×× +=∑ ,
[ ] TZCXCC TSTSTS
rrr 0 ,, −= [ ] TZTYTXTT CNCNCNCN
FFFF ,,,=
[ ] TXSS TSTS
rr 00,−= [ ] TZSXSS TSTSTS
FFF 0 ,,=
The term CTS,CN in the above equation (Eq. 1) is the transformation matrix that translates
the thrust from the CNAPS frame to the thrust stand frame. This is the only term in the
equation that changes for each of the three cases below. All equations and references
made below refer to the thrust stand reference frame unless otherwise noted. For clarity,
the vectrix TTSℑ has been omitted from equations.
Case 1: CNAPS X-axis aligned with thrust stand Z-axis
CNAPS will typically be oriented in the nominal orientation, i.e. with the CNAPS X-axis
aligned with the Z-axis of the thrust stand. All of the tests for characterizing the
performance of the propulsion system, as outlined in [19], will be conducted in this
configuration.
56
Figure 31 CNAPS orientation for Case 1
The transformation matrix CTS,CN is applied to the thrust vector, FT, to determine its
components in the thrust stand frame.
−−
=001
100
010
,CNTSC
Equation 2 is as follows, using the general Eq. 1:
( )
( )
( )
=
+
−−
−
=
−+
−−
−−=∑
0
0
0
0
0
0
0
0
0
00
00
000
001
100
010
00
0
00
1,,
,,
,,,,
,,
1,
1,
,
,
,
,
,
,
,,
,
ZSXS
ZTXC
XTXCYTZC
ZTZC
ZS
XS
XS
XS
ZT
YT
XT
XC
XCZC
ZC
O
Fr
Fr
FrFr
Fr
F
F
r
r
F
F
F
r
rr
r
M
The moment produced by the misaligned thrust vector will be made up of two
components: the X and Z component as seen in Figure 31. The force measured at the
sensor can be converted to the corresponding moment using the known distance between
the sensor and the pivot point, rS. Therefore, in the above equation, only the two
components of the thrust vector FT,X and FT,Y are unknown. FS,Z (1) is the force detected by
the sensor, where the subscript 1 indicates the measurement for Case 1.
57
Figure 32 Nozzle placement with respect to centre of CNAPS
Case 2: CNAPS Y-axis aligned with thrust stand Z-axis
With CNAPS oriented in this configuration, the sensor can detect forces again in the XZ
plane. The H-frame is used to fasten CNAPS to the thrust stand. In this orientation, the
moment arm of the thrust component in the Z-direction is minimized to reduce its affect
on the resulting moment, and thus force measured by the sensor. At the same time, the
moment arm for the off-axis thrust component in the X-direction is sufficiently
maximized for the sensor to detect forces in that direction, even for the minimum thrust
case where a single thruster produces 5 mN of thrust.
58
Figure 33 CNAPS orientation for Case 2
The transformation matrix for Case 2, and the corresponding moment equation are as
follows:
−−
−=
010
100
001
,CNTSC
Eq. 3:
( )
( )
( )
=
+
−+
−
=
−+
−−
−
−−=∑
0
0
0
0
0
0
0
0
0
00
00
000
010
100
001
00
0
00
2,,
,,
,,,,
,,
2,
2,
,
,
,
,
,
,
,,
,
ZSXS
ZTXC
YTXCXTZC
ZTZC
ZS
XS
XS
XS
ZT
YT
XT
XC
XCZC
ZC
O
Fr
Fr
FrFr
Fr
F
F
r
r
F
F
F
r
rr
r
M
As in Case 1, the moment produced by the thrust vector consists of X and Z components.
However, the moment arms in the X and Z axes differ from those in Case 1 due to the
orientation of CNAPS. As shown in Figure 32, the distance between the geometric centre
of the four-nozzles and an individual nozzle is relatively small when compared to the
moment arms of the thrust components. Thus, in the case where only 1 nozzle is being
used, the corresponding change in the measured force is small yet noticeable since it is
well within the range of the sensor.
59
Case 3: CNAPS Z-axis aligned with thrust stand Z-axis
The H-frame is used to attach CNAPS to the multiplier arm such that its Z-axis is aligned
with the thrust stand’s Z-axis. In this setup, the sensor is only capable of measuring
forces in the XZ plane of the CNAPS reference frame, with forces in the Y-axis of the
same frame being cancelled due to the pivot design. The moment produced by the along
axis thrust component in the X-direction is minimized due to the shorter moment arm,
whereas the effect of the off-axis component in the Z-direction is maximized due to the
larger moment arm.
Figure 34 CNAPS orientation for Case 3
The transformation matrix for Case 3, and the corresponding moment equation are as
follows:
−−
=100
010
001
,CNTSC
60
Eq. 4:
( )
( )
( )
=
+
−−
−
=
−+
−−
−−=∑
0
0
0
0
0
0
0
0
0
00
00
000
100
010
001
00
0
00
3,,
,,
,,,,
,,
3,
3,
,
,
,
,
,
,
,,
,
ZSXS
YTXC
ZTXCXTZC
YTZC
ZS
XS
XS
XS
ZT
YT
XT
XC
XCZC
ZC
O
Fr
Fr
FrFr
Fr
F
F
r
r
F
F
F
r
rr
r
M
6.1. Thrust Components
In order to determine the individual components of the thrust vector, equations 2-4 can be
solved.
( )
( )
( )
−−−
=
−
−
3
2
1
,,
,,
,,
,
,
,
,,
,,
,,
0
0
0
ZSXS
ZSXS
ZSXS
ZT
YT
XT
XCZC
XCZC
ZCXC
Fr
Fr
Fr
F
F
F
rr
rr
rr
7. Calibration
The thrust stand and force sensor needed to be calibrated before proceeding with testing
CNAPS. This included determining an appropriate scaling factor for the charge amplifier
that converts the sensor’s charge to a voltage. As well, noise sources had to be identified
and isolated to reduce their effects on the results of the tests. Due to the high level of
accuracy and resolution required for the results of CNAPS’ performance tests, all effects
of the operating environment and equipment would need to be accounted for.
7.1. Setup
Calibrating the thrust stand involved comparing the output from the sensor under the
influence of a known load placed on stand. To do this, the thrust stand was setup exactly
61
how it would be during testing CNAPS; it was placed inside SFL’s vacuum chamber and
the appropriate cables and connectors were used, even if not required, to capture their
influence on the output. The only difference between the setup for calibration and that for
actual test was that the vacuum chamber door was left open to allow for modifications to
be made to the thrust stand during calibration.
A circular disk with a conical cut-out in its centre, was fastened to the end of the arm
intended for CNAPS as shown in Figure 35. The conical cut-out was to ensure that the
gravitational force of the mass placed in the disk passed through its centre, making it
easier to determine its moment arm with respect to the pivot. The arm was balanced by
placing counterweights on its other end so that preloading on the sensor was minimized.
Figure 35 Setup of thrust stand inside SFL vacuum chamber
7.2. Data Collection
An industrial grade data acquisition unit was used to measure the change in force when a
mass was placed on the thrust stand. Although the calibration test itself did not require a
high temporal resolution for measurement, the DAQ was included as part of the test setup
because it would be used to collect data during performance testing of CNAPS. Data
62
collected by the DAQ was stored as a ‘.csv’ file containing the voltage output by the
charge amplifier, which was later post-processed.
The detection of the force produced by placing a mass on the thrust stand was represented
by a change in voltage. Collected raw data was plotted to identify the point at which the
mass was placed, which is made evident by a relatively large change in voltage. By
measuring the average voltage before and after the mass is placed to filter out noise, and
subtracting one from the other, the voltage difference corresponding to the mass was
derived. By multiplying the change in voltage with the scaling factor, and using the
moment arm ratio between the disk and sensor, the equivalent mass in the disk was
determined.
Figure 36 Sample plot of measured force showing data extraction range
For masses of 1 g or greater, the voltage difference was easily identifiable for all scaling
factors on the charge amplifier. This was not true for masses of 100 mg or less. It was
noticed that for these smaller masses, a scaling factor of 1 made it difficult to extract the
change in voltage produced by adding the mass. Proof of this is shown in the following
63
section. By using a smaller scaling factor, the charge amplifier was able to detect the
difference in charge of the force sensor to a finer level.
Figure 37 Sample plot of measured force for small mass and scaling factor
On the other hand, using a small scaling factor meant that the charge amplifier was even
detecting low-level background noise and vibrations which was affecting the accuracy of
the measurement. Therefore, for small scaling factors, the ‘instantaneous’ change in
voltage was extracted; i.e. a smaller range of data before and after adding the mass was
averaged.
7.3. Results
The detailed procedures followed for calibrating the thrust stand may be found in [21]. In
summary, various precision masses representing different thrust profiles and fuel
quantities of CNAPS, were placed in the disk at the end of the multiplier arm. The change
in sensor output was recorded, from which the mass in the disk was calculated using the
known moment arm ratio. The measured mass was compared to that specified in the
calibration certificate of the precision masses. Different scaling factors on the charge
64
amplifier were tried, to determine the best one to be used for a particular force to be
measured.
One of biggest issues that arose in the beginning of the tests was the drift in the output of
the sensor. Upon researching this matter, it was found that a manufacturer-recommended
procedure was to be followed while cleaning all cable ends and connectors. After
cleaning all cables and connectors, the drift was seen to reduce to an acceptable level.
Mass vs Scaling Facor
0.00%
0.50%
1.00%
1.50%
2.00%
2.50%
3.00%
0.05 0.1 1 10 30
Mass (g)
Err
or (
%)
Scale 0.3Scale 0.5
Scale 1
Figure 38 Plot showing relationship between measurement error and scaling factor
The general trend observed was that a smaller scaling factor produces more accurate
measurements (to within 1%) especially for smaller masses, up to 1 g. While this is also
true for masses greater than 1 g, the maximum measuring range of the charge amplifier
limits the maximum force that can be measured by the thrust stand. Even for a mass of 5
g, the small shock produced when placing the mass onto the disk was seen to exceed the
measuring capacity of the charge amplifier for the given measuring range.
In the case of testing CNAPS on the thrust stand, the choice of scaling factor will depend
largely on the test being performed. For measuring the thrust profile over a short
65
duration, as well as determining the specific impulse, a scaling factor in the range of 0.3-
0.5 would be suitable for measuring the thrust force and the change in mass of CNAPS
due to venting of SF6. Tests requiring measurement of the thrust profile over periods
lasting more than 15 seconds would require a scaling factor in the range of 0.5 – 1 due to
the larger mass change observed from venting SF6.
The charge amplifier has the added benefit of being reset before each test, so that only the
relative change in mass of fuel for that particular test can be measured. Resetting the
charge amplifier for each test removes the effect of the drift, characteristic of force
transducers, from the final result. While short duration tests may not be influenced very
much, the drift rate may affect the accuracy of the result for long duration thrusts.
8. Recommendations and Future Work
The results from the calibration tests of the thrust stand show that it is capable of
measuring thrust forces up to a resolution of 0.5 mN with an error margin of less than
3%. This satisfies the requirements for the thrust stand. Although not shown in the bar
graph (Figure 38) tests were also conducted with a 10 mg mass. However, detection of
the mass being placed on the thrust stand was overshadowed by noise and drift picked up
by the sensor. This made it difficult to extract the change in measured force from the
collected data, resulting in large errors. This is not of concern since this translates to a
measurement resolution of 0.1 mN for the sensor, which is only desired. The required
resolution is of 0.5 mN and is detectable by the sensor.
The next step will be to proceed with testing a fully assembled CNAPS. Down the line, as
the formation flying algorithm approaches maturity, the tolerances on the precise
knowledge of CNAPS’ thrust vector might become smaller, requiring the thrust stand to
be modified and recalibrated for measuring forces to a higher resolution and accuracy.
This can be accomplished by either relaxing the dimensional requirements of the thrust
stand, or by recalibrating it for a smaller scaling factor.
66
Chapter 3
CNAPS Pressure Regulator
Valve Thermal Analysis and
Testing
1. Introduction
CNAPS shares a similar architecture to NANOPS. This includes regulating the pressure
from the vapour pressure in the primary tank to a lower pressure for finer thrust control.
Volume constraints on NANOPS forced the removal of the secondary volume, causing
the regulator valve to be connected directly with the thrust valve [23]. In CNAPS, the
secondary volume serves to stabilize the flow and maintain a constant pressure to feed the
thrust valves.
Previous experience on NANOPS with the regulator valve showed that the performance
of the valve degrades during extended on-times of and flow rates through the valve. The
thermodynamic effect of venting SF6 to vacuum for extended periods causes the valve
temperature to decrease below its operational limits, affecting its performance. Moreover,
low thermal contact between the valve and surrounding structure limits the amount of
energy entering the valve. The regulator valve to be used on CNAPS consists of a
67
cylindrical body with flanges on either side. It is fitted to the CNAPS structure through
four holes on it flanges, with screw-spacer pairs.
Degradation in valve performance is of concern for CNAPS because the CanX-4/-5
mission requires extended on-times for the regulator valve during orbit reconfiguration
manoeuvres. In addition, past experience has shown that when the regulator valve’s
performance diminishes, the flow rate through the valve is affected. Because the
formation flying mission requires precise knowledge and control over the thrust profile, a
change in the flow rate, and hence the secondary tank pressure, was not desired.
To better understand the valve’s performance with respect to temperature variation and
on-times, a study was conducted. This involved creating an analytical model to analyze
the thermal behaviour of the valve and then comparing the predicted temperatures to
actual values obtained from a test.
2. Design
The current CNAPS design consists of a tray structure housing all the propulsion
components. The regulator valve (RV) is connected in line between the primary and
secondary volumes. It is fastened to the tray cover through four screw holes on its
flanges. The valve’s inlet and exit ports are connected to the surrounding CNAPS
structure through tubing.
68
Figure 39 Regulator valve fastened to CNAPS structure
The regulator valve consists of a central plunger pushed against a rubber seal with a
spring. When the valve is actuated, the plunger moves, allowing fluid to flow through the
valve. The housing of the RV is made of stainless steel. A combination of frequency and
duty cycle determines the mass flow rate of SF6 through the valve.
Figure 40 Regulator valve cross section (Courtesy Mike Borla, The Lee Co.)
3. Analysis
Before proceeding with testing, an analytical model of the valve was created. The
purpose of this was to know what temperature variations to expect during the test. As per
standard SFL policy, the analytical model is required to predict temperatures to within 5
C of measured temperature. This section first briefly goes over the relevant theory
concerning the test, followed by a mathematical model of the valve along with the
expected temperature change.
69
3.1. Relevant Theory
To analyze the thermal behaviour of the valve, it was necessary to know the
characteristics of the flow passing through it. Once this was determined, heat transfer
methods could be employed to predict the temperature variation with respect to time. The
analysis was categorized into three stages, as follows. Beginning with isentropic relations
to determine the mass flow rate through the valve, the thermodynamics of the phase
change of SF6 from liquid to vapour, and finally the heat transfer, both within, and to and
from the valve, were calculated. The governing equations and theory for each are
described below.
3.1.1. Isentropic Flow
Most cold gas propulsion systems use isentropic relations to predict performance [23].
Given the nozzle dimensions and pressure of SF6 stored in the secondary volume, the
mass flow rate can be calculated using the equation shown below. For given stagnation
pressure, pO, stagnation temperature, TO, Mach number, M, specific heat ratio of the
fluid, γ, and throat diameter A*, the mass flow, rate is determined by:
( )( )12
1
0
0*
12 −
+
+=
γγ
γγ
RT
pAmɺ
The SF6 stored in the secondary volume is assumed to be at its stagnation temperature
and pressure, since the fluid inside does not experience too much variation during normal
CNAPS operations. Therefore, using the throat area of a single nozzle downstream, the
maximum mass flow rate through the regulator valve was calculated to be approximately
0.026 g/s. In the case all four thrusters are turned on, the flow rate through the valve is
four times this value, which is 0.1 g/s.
70
Typical operations of CNAPS will involve three thrusters being used at a given time to
reduce the effects of momentum build-up [24]. However the above mass flow rate
assumes a worst case scenario of all four thrusters being used simultaneously.
3.1.2. Thermodynamics
Nominal Operations
Ideally, CNAPS will run off the secondary volume, where the pressure of SF6 is regulated
down to 6.9 bar from its vapour pressure, which is temperature-dependent, in the primary
volume. Therefore, the initial thermodynamic analysis was done assuming SF6 at 6.9 bar
in the secondary volume.
For the purpose of evaluating how much energy is consumed by SF6 when it expands
from vapour pressure to secondary volume pressure, enthalpy was considered since it
encompasses the internal energy at the molecular level, and the work done by the energy
in the flow. Property tables of empirical data from [25] were consulted for the enthalpy
values of SF6.
In the interest of being conservative, it was assumed that SF6 in the primary volume has
the lowest possible enthalpy for the suggested CNAPS temperature range, and is
expanded through the RV to a pressure of 6.9 bar in the secondary volume. This gives the
largest change in enthalpy that will most likely occur. For a nominal temperature range of
10 to 40 C as specified in the thermal requirements of CNAPS [26], the lowest possible
enthalpy value of SF6 is 210 kJ/kg, at a temperature of 10 C. Moving away from the
saturation curve, a pressure of 6.9 bar in the secondary storage tank, at a temperature of
10 C, corresponds to an enthalpy of 296 kJ/kg for SF6 vapour. On the other end of the
temperature range, the enthalpy values for liquid and vapour were seen to be 245 kJ/kg
and 317 kJ/kg respectively at 40 C. For the case where CNAPS temperature is 10 C, a
larger enthalpy change of 86 kJ/kg occurs through the valve, and hence this case was
chosen for the analysis in the interest of being conservative.
71
The total energy consumed by SF6 as it evaporates during pressure regulation is the
product of the mass flow rate and the net enthalpy change. Using the mass flow rate
determined in an earlier section, the net power loss from the valve due to SF6 phase
change is shown below.
( )( )sJE
gJsgE
hmE
out
out
out
6.8
861.0
=
=
∆=
ɺ
ɺ
ɺɺ
Extended Operations
During reconfiguration manoeuvres of CanX-4 and CanX-5, an extended duration of
thrust is required to provide the necessary ∆V. The secondary storage tank has a capacity
an order of magnitude less than the primary tank. Due to this, there is the likelihood that
the smaller tank will be partially emptied during long duration thrusts, thereby causing
the primary tank to act as the SF6 supply for the thrust. In this case, SF6 at a much higher
pressure in the primary volume will undergo expansion through the RV, leading to a
larger enthalpy change in the flow. The amount of energy absorbed by SF6 during the
expansion will be taken from the RV.
The likelihood of CNAPS running directly from the primary tank is dependent on many
variables, including the formation flying algorithm currently being developed. Testing
will confirm the possibility of this scenario occurring. However, taking a conservative
approach, this case was looked at from a thermal perspective.
Again, the largest possible enthalpy change was considered during SF6 phase change, as a
way of encompassing all cases with smaller changes in enthalpy. It was assumed that SF6
stored in the primary tank is at vapour pressure. The largest enthalpy change is noticed at
the lower end of the temperature range of CNAPS, at 10 C, with SF6 at a vapour pressure
of 16 bar, corresponding to an enthalpy value of 209 kJ/kg. When it is vented to the
vacuum of space, SF6 in its vapour phase has an enthalpy of 301 kJ/kg, for an adiabatic
72
expansion process. This gives a net enthalpy change of 92 kJ/kg. The energy loss from
the valve is determined using the mass flow rate calculated earlier.
( )( )sJE
gJsgE
hmE
out
out
out
2.9
921.0
=
=
∆=
ɺ
ɺ
ɺɺ
3.1.3. Heat Transfer
The regulator valve is fastened to the CNAPS tray cover by its flanges, with four pairs of
stainless steel screws and aluminum spacers. In the axial direction, the inlet and exit ports
of the valve connect to PEEK tubing. This limits the conduction path to the screws and
spacers between the RV and the tray cover.
While creating the model, thermal paths were modelled in terms of parallel resistors in a
circuit. The total resistance of the four screw-spacer pairs was determined by summing
the resistance of each pair, which included the following: contact resistance between the
screw, spacer and tray respectively; resistance through the screw and spacer, and the
valve flange. While resistance through the screw and spacer could be determined easily
based on dimensions and material properties, the biggest uncertainty was the contact
resistance values. Empirical values were used in the analysis, which were obtained from
[27].
73
Figure 41 Equivalent circuit diagram for thermal resistsance
An energy balance approach was employed to determine the temperature change in the
valve during thrusts. By treating the valve as a control volume, the difference between the
rates of energy entering and leaving the system was equated to the change in the rate of
internal energy of the system.
Figure 42 Control volume around regulator valve
( ) ( ) storedoutflowSFinconductionflowSF
storedoutin
EEEE
EEE
ɺɺɺɺ
ɺɺɺ
=−+
=−
66
Energy entering the system is the conduction of heat from the tray, and the energy
brought in by SF6, given by its enthalpy value. Energy leaving the valve was modelled as
the enthalpy of SF6 flowing out. The conduction path between the tray cover and the
74
valve was modelled like an electrical circuit, with nodes representing the valve and cover.
The screws and spacers were modelled as resistors in the circuit.
Thermal resistance of an object is a function of its dimensions and material properties.
Typically, conductive resistance is expressed as:
kA
lR =
k = thermal conductivity of material
l = dimension of object parallel to heat flow path
A = cross sectional area of object perpendicular to heat flow path
For radial systems, such as tubes and hollow cylinders, resistance in the radial direction is
expressed as:
Lk
rr
Rπ2
ln1
2
=
r1 = inner diameter
r2 = outer diameter
k = thermal conductivity of material
L = length of cylinder that is being considered
The resistance from the valve body to the tray cover was modelled as a chain of resistors
in series and parallel in an electrical circuit. Resistance in series can simply be added
together. The total resistance for parallel resistors can be found using the equation shown
below.
∑=
=N
i itotal RR 1
11
75
i = resistance through a particular component or interface
N = total number of resistors in parallel
3.2. Thermal Modelling
During typical operations of CNAPS, the secondary volume acts as the supply tank for
SF6. Thus, the regulator valve adjusts the pressure from the vapour pressure in the
primary tank, to a lower pressure in the secondary tank. A length of PEEK tubing exists
between the valve and primary and secondary tanks respectively. This causes SF6 to
undergo expansion within the tube downstream of the valve. However, since the length
and diameter of the tube had still not been selected, a conservative approach was taken
and the assumption was made that SF6 evaporates within the valve body, while passing
through the orifice controlled by the plunger and seal. This approach, although not very
realistic, captures the limiting temperature change experienced by the valve.
Different cases were examined using the model, as outlined below.
Case 1: Nominal Operations
Nominal operations of CNAPS represent the case where the secondary tank acts as SF6
source for the thrusters. Typical short duration thrusts, lasting under 15 seconds fall under
this category. During nominal operations, the mass flow rate of SF6 through the valve is
approximately 0.1 g/s, assuming all four thrusters are in use. Using conservation of
energy principles, the net change in valve temperature was determined.
Initial results based on the analytical model showed that the valve temperature fell
significantly below its operational limits (-29 to +177 C) even for short duration thrusts
(Figure 43). The analysis was redone by increasing the thermal coupling between the
valve and tray cover, assuming the use of a gap filler between the two (Figure 44). Initial
76
temperature assumed for the analysis (15 C) was the lowest operational temperature
predicted for CNAPS based on preliminary thermal analysis [20].
Temperature Profile for CNAPS RV - Nominal operatio ns without gap filler
-100
-80
-60
-40
-20
0
20
0 20 40 60 80 100 120
Time (s)
Tem
pera
ture
(C
)
No heater
Figure 43 Temperature profile of valve without gap filler
Temperature Profile for CNAPS RV - Nominal Operatio ns
0
2
4
6
8
10
12
14
16
0 10 20 30 40 50 60 70 80 90 100 110
Time (s)
Tem
pera
ture
(C
)
No heater
1 W heater
Figure 44 Temperature profile of RV for nominal operations with gap filler in use
77
The addition of the gap filler in seen to improve the thermal coupling between the valve
and tray cover. A steady state temperature of 2 C is reached even for thrust times of up to
100 s. The graph also shows the effect of adding a 1 W patch heater to the valve; in this
case the steady state temperature reached is approximately 3.5 C.
Case 2: Extended Operations
Extended operations of CNAPS represent the extended on-times of the regulator valve.
This usually happens during reconfiguration manoeuvres when a large ∆V is required and
the thrusters are turned on for durations of up to 70 seconds. In this case, the SF6 in the
secondary volume gets exhausted and the regulator valve begins to allow SF6 to flow
from the primary to the secondary tank to maintain the required pressure of 6.9 bar. This
extended on-time coupled with the larger mass flow rate passing through the regulator
valve causes a greater temperature change as seen in Figure 45.
Temperature Profile for CNAPS RV - Extended Operati ons
-25
-20
-15
-10
-5
0
5
10
15
20
0 10 20 30 40 50 60 70 80 90 100 110
Time (s)
Tem
pera
ture
(C
)
No heater
Figure 45 Temperature profile of RV for extended operations with gap filler in use
The analysis showed that the steady state temperature of approximately -21 C is
acceptable since it is within the operational temperature range of the valve (-29 to +177
C), even without a heater being used.
78
Case 3: Regulator valve used for testing
Rather than testing the flight valve directly, it was decided at the managerial level that a
spare valve would first be tested to note temperature variation over extended operations.
The valve selected is identical in operation to the CNAPS flight valve with the following
two exceptions: the valve is rated to 500 psi, as opposed to the 1000 psi rating for the
flight valve; and, the exterior of the valve does not have flanges. Therefore, a separate
model was created for this valve.
Since the test valve does not have flanges, it would require a clamping mechanism to
fasten it to the tray cover during tests. Therefore, a bracket made of a non-conductive
material was chosen for this purpose, and was included in the analytical mode. The added
resistance due to the bracket allowed for a more conservative estimate of the valve
temperature. The initial temperature chosen for the analysis was 22 C, to represent the
room temperature conditions under which the bench test would be carried out.
Temperature at RV Exit Port - Predicted Model 2
14
15
16
17
18
19
20
21
22
23
0 20 40 60 80 100 120
Time (s)
Tem
pera
ture
(C
)
Figure 46 Temperature profile of test RV with gap filler in use
The analytical model predicts the temperature of the valve exit port closest to the main
valve body. This was done so that the model could be compared to the temperatures
79
measured at that point, during tests. It is believed that the temperature variation between
the valve orifice and the point at which temperature is measured will be very small
because of the relative proximity of the two (Figure 40).
4. Bench Testing
4.1. Setup
The test setup included a semi-assembled CNAPS structure in order to test the valve in
conditions representative of the actual flight hardware. The CNAPS tray and side panels
were assembled, with a single primary tank fastened to the inside of the panels. Only one
tank was used for the purpose of testing the valve since its capacity was more than
sufficient to carry out the required tests. PEEK tubing was used to connect the valve to
the primary tank, with a pressure transducer connected in series. The bottom tray cover
was left out from the assembly to allow access to the valve and temperature sensors
during tests. SF6 was vented from the valve’s exit port, directly to the atmosphere.
Figure 47 Setup for thermal testing of RV
80
The regulator valve used for testing, although similar in function and operation to the
CNAPS valve, does not have flanges on its outer body. Therefore, a bracket was designed
to hold the valve against the tray cover during tests as seen in Figure 48. Delrin was
chosen as the material for this bracket due to its low thermal capacitance and availability.
The bracket was made from a block of Delrin, with a groove machined on its side to sit
the valve in. Screw holes on either side of the groove were made to align with the already
existing holes on the tray cover.
Figure 48 Test valve fastened to CNAPS with Delrin block (front view)
A commercially available thermal gap filler was inserted in between the valve body and
the tray cover as a way of increasing the conduction path between the two. Since the test
valve does not have flanges, and is held by a thermally insulated bracket, the gap filler
layer provides the primary path for heat flow from the tray cover to the valve.
4.2. Data collection
Thermistors were chosen as a means of measuring temperature along various points on
CNAPS during tests. These small bead-like sensors provide a convenient way of
attaching them to tubes and unions. Thermistors have the property of having their
resistance changed as a function of the surrounding temperature. However, the lab data
81
acquisition unit (DAQ) is only capable of measuring voltage differences. This required
creating a breadboard circuit with a resistor and capacitor, to measure the voltage drop
across the thermistor. A schematic of it is shown in Figure 49. The change in resistance
across the thermistor was then determined from the corresponding voltage output from
the DAQ.
Thermistor beads were placed at various points along the test setup. Since the valve body
is cylindrical and sits pressed against the gap filler layer and tray cover on one side, and
the Delrin bracket on the other, it wasn’t possible to have a sensor on the valve body
itself. Moreover, the body is a cylinder of much larger diameter than the exit and inlet
ports. So, it was believed that measuring the temperature at the exit port of the valve
would provide a better indication of the temperature inside the main body due to its
proximity to the orifice (Figure 40). Two thermistors were also placed on the tray cover,
one on each side directly over the valve to note any change in temperature. Another
thermistor was positioned at the end of the exit. Since the thermistors are circular beads,
placing them against a surface only provided a point contact between them and the
surface. So gap fillers were used to ensure that the temperature of the surface was
properly conducted to the thermistor. Kapton tape was used to secure the sensor and gap
filler to the surface of interest. Prior to testing the valve, the thermistors were connected
to the DAQ and the measured output voltage and corresponding resistance was compared
to empirical values provided by the manufacturer to verify their performance.
Output from the DAQ was logged on a computer and stored in an Excel file. Pressure
data was collected from the transducer via a circuit board designed for use on NANOPS.
The same board was also used to send commands from the computer to actuate the
regulator valve.
82
Figure 49 Schematic of breadboard circuit
The resistance of the thermistor is a function of the fixed resistor, and the input and
output voltage. The temperature measured by the thermistor was calculated from this
resistance value using the Steinhart-Hart equation [28].
outin
outNthermistor VV
RVR
−=,
( ) ( )
8
4
3
3
1090.9
1037.2
1040.1
lnln1
−
−
−
×=
×=
×=
++=
c
b
a
RcRbaT
83
4.3. Tests
Test procedures included initializing the DAQ and making sure the thermistors
temperatures were stabilized before beginning the test. During tests, the output form the
thermistors was logged via the DAQ, while pressure was measured via the NANOPS
electronics board with the pressure transducer connected in line with the primary tank
and valve. A command was sent manually to actuate the regulator valve. Starting with 10
seconds, valve on-times were increased in 10 second intervals in subsequent tests, up to
an on-time of 60 seconds. A wait time of 30 minutes between tests was included in the
test procedure to ensure that the components and temperature sensors returned to room
temperature before beginning each test. Detailed test procedures can be found in the test
plan document [29].
Figure 50 Test valve with regions of interest identified
Figure 51 show the lowest temperatures measured on the valve for increasing on-times.
One discrepancy to note is the slightly higher temperature observed for the 60 s on-time
test, compared to the 50 s test. The likely cause of this is that the pressure inside the
primary tank decreased below the vapour pressure, therefore causing the SF6 in liquid
phase inside the tank to boil off. Rather than having SF6 undergo a phase change from
liquid to vapour through the valve, it would merely be undergoing an expansion.
84
RV exit port temperature vs on-time
16
16.5
17
17.5
18
18.5
19
19.5
20
10 20 30 40 50 60
Valve on-time (s)
Tem
pera
ture
(C
)
Figure 51 Minimum temperatures measured at RV exit port for different on-times
Table 4 Predicted and measured minimum temperatures at valve exit port
On-time Measured Predicted
10 19.4 17.9
20 18.2 16.4
30 17.5 15.8
40 17.2 15.6
50 17.1 15.5
60 17.3 15.5
Min Temperature ('C)
The temperature at the valve exit port seemed to reach a steady state value of
approximately 17 C with increasing valve on-times. Although there were thermistors
placed at other points on the test setup, the temperature variation seen at those points was
negligible compared to that seen at the valve exit port. The thermistor at the end of the
valve’s port was placed as a check to ensure that SF6 indeed was changing phase while
passing through the valve.
85
Upon comparing the measured temperatures at the valve exit port to those based on the
mathematical model, it is seen that the predicted numbers are off by up to 2 C (Table 4).
This discrepancy between the predicted and measured temperatures may be explained by
the uncertainties related to contact resistance and the limited knowledge of the interior of
the valve for modelling purposes. Overall the model is seen to capture the thermal effects
on the valve to an acceptable level; standard SFL policy dictates that the analytical model
be able to predict temperatures to within 5 C of those measured.
5. Conclusion
It is worth noting that the conditions for the bench test of the valve were fairly
conservative. The valve was more thermally isolated than it would be in flight-like
conditions. This means that the CNAPS flight valve will likely not see the temperatures
predicted by the analytical models. In practice, the flight hardware will consist of the
valve in between the primary and secondary tanks. Since the secondary tank pressure is
always maintained much higher than atmospheric pressure throughout the mission, the
SF6 phase change energy, and hence valve temperature variation, will be much smaller
than that experienced during bench testing in the laboratory and predicted by the
numerical models.
To further validate the models for the valves, additional tests will be carried out with the
test valve in vacuum conditions, and finally with a spare flight valve itself. Results from
the bench tests showed that the enthalpy change of SF6 was the dominating cause of the
temperature variation of the valve; convection seemed to have little influence on the
valve temperature. Based on this, it is expected that the results from the tests in vacuum
conditions will not be very different from those obtained from bench tests. Future tests
would need to be performed with a fully assembled and fuelled CNAPS to represent
flight-like conditions. The ambient temperature may be varied to correspond to the range
of temperatures experienced by CNAPS during on-orbit operations to examine the
influence of temperature on the enthalpy of stored SF6.
86
Conclusion
The satellite’s deployment speed form the XPOD was successfully measured by the
ground support equipment developed for the test. Having the GTV attached to the
satellite was seen to have little influence on its deployment speed compared to its
theoretical performance. The GTV successfully integrated with the satellite without
interfering the XPOD’s release and deployment mechanism. Moreover, from all the tests
that were conducted, both with and without the XPOD armed, the GTV did not tip over
or collide with the XPOD or XPOD Mount structure. Therefore, having proved the
reliability of the GTV with tests done with the mass dummy, the next step would be to
test a fully assembled spare satellite structure before moving onto testing a satellite ready
for flight.
Calibration tests of the thrust stand show that it is capable of measuring the minimum
thrust produced by CNAPS, which is expected to be 5 mN. Also demonstrated during the
calibration was the thrust stand’s ability to detect forces with a resolution of 0.5 mN,
which meets the tolerance requirements imposed on the CNAPS’ thrust magnitude. The
thrust stand was also shown to be capable of determining the misalignment of the thrust
vector by measuring components of the thrust force along each axis. A separate adapter
was designed to orient CNAPS in different configurations to measure all components of
thrust. Future work includes testing the performance of CNAPS on the thrust stand, in
high vacuum conditions so that any issues with the design or assembly of the propulsion
system can be identified early on in the integration and test phase.
Tests done with the regulator valve have shown that the average valve temperature
remains within its operational limits even during extended valve on-times. The test
conditions were fairly conservative. Larger pressure differentials and thus mass flow
rates, were used to gain confidence in the valve’s performance over extended on-times,
and to validate the analytical models used to predict temperature variation. Tasks for the
87
near future include testing a spare flight valve, first in atmospheric conditions, and then in
high vacuum conditions for ambient temperatures spanning the operational temperature
range of CNAPS.
88
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