RESEARCH PAPER
Vibration isolation system for cryocoolers on the
XRISM
Susumu Yasuda ,a,* Yoh Takei ,a Atsushi Okamoto,a Keiichi Yanagase ,a
Kazuki Watanabe,b and Seiji Yoshidac
a
ABSTRACT.
Japan Aerospace Exploration Agency, Tsukuba, Japan
b
WEL Research, Ichihara, Japan
c
Sumitomo Heavy Industries, Ltd., Niihama, Japan
We discuss the design and performance of the vibration isolation system (VIS) used
for the Resolve microcalorimeter on the X-Ray Imaging and Spectroscopy Mission
(XRISM) spacecraft. Resolve makes high spectral resolution observations of
celestial X-rays. XRISM was conceived as a recovery mission following the
short-lived Hitomi mission. As such, the design of Resolve is similar to that of the
soft X-ray spectrometer (SXS). A VIS was initially developed for the Hitomi SXS after
discovering that vibrations from the cryocooler degraded the sensor’s performance.
However, the VIS for Resolve was completely redesigned to include a hold-andrelease mechanism to avoid damage to the compressor during launch or
ground-based sine-vibe tests. The XRISM spacecraft was successfully launched
in September 2023, confirming the functionality of the VIS while in orbit.
© The Authors. Published by SPIE under a Creative Commons Attribution 4.0 International License.
Distribution or reproduction of this work in whole or in part requires full attribution of the original
publication, including its DOI. [DOI: 10.1117/1.JATIS.11.4.042004]
Keywords: X-ray observation; microcalorimeter; cryocooler; vibration isolation
Paper 24213SS received Dec. 13, 2024; revised Feb. 13, 2025; accepted Feb. 21, 2025; published Mar.
12, 2025.
1 Introduction
Cryogenic techniques are often used in space missions that require extremely high sensitivity or
low noise measurements. Cryogenic instruments have revealed an unprecedented view of the
universe at various wavelengths. Examples include COBE,1 WMAP,2 and Planck3 for observing
the cosmic microwave background; AKARI,4 Spitzer,5 Herschel,6 and JWST7 for observing the
universe in the infrared; and Hitomi8 for X-rays.
Cryogens, particularly liquid (or superfluid) helium, have been used for a long time in space
missions. In addition, cryocoolers have recently been used to reduce the heat load on the cryogen
(e.g., AKARI, Hitomi, and XRISM), or even as a primary cooling method (e.g., JWST).
Furthermore, many planned missions (e.g., LiteBIRD,9 Athena X-IFU,10 Line Emission
Mapper,11 Probe Far-Infrared Mission for Astrophysics,12 and Black Hole Explorer13) will operate without cryogen, instead relying on cryocoolers, so that the mission lifetime is not limited by
the amount of cryogen.
Cryocoolers [e.g., Stirling coolers (ST), pulse-tube coolers, and Joule-Thomson coolers
(JT)] utilize repeated cycles of compression and expansion of the working gas for cooling.
The mechanical movement of compressors (CMPs) and other parts, along with the flowing working gas, creates vibration, often disturbing highly sensitive onboard instruments. Mitigating these
vibrations is therefore important.
*Address all correspondence to Susumu Yasuda, yasuda.susumu@jaxa.jp
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This paper presents the design and performance of a vibration isolation system (VIS) for the
Resolve instrument14 on the X-ray imaging and spectroscopy mission (XRISM).15 Resolve is a
microcalorimeter-based instrument enabling observations of celestial X-rays with high spectral
resolution (E∕ΔE ∼ 1000, or <7 eV full-width half-maximum at 6 keV).
As XRISM was a recovery mission for the short-lived Hitomi, the design of Resolve is
fundamentally similar to the soft X-ray spectrometer (SXS),16 the microcalorimeter instrument
onboard Hitomi. Thus, the design of Resolve’s cooling chain is the same as that of SXS. Details
can be found in Refs. 17 and 18. The cross-section of the Dewar (DWR) and its cryogenic architecture are shown in Figs. 1 and 2, respectively.
The detector is installed inside the helium DWR, a vacuum vessel containing a helium tank.
A three-stage adiabatic demagnetization refrigerator (i.e., ADR19) achieves the operating temperature of the detector, 50 mK. The assembly containing the detector and the ADR (called the
calorimeter spectrometer insert; CSI) is mounted on a 39 L helium tank and cooled to around
1 K by superfluid helium. There are multiple radiation shields between the helium tank and the
DWR outer shell [i.e., the DWR main shell (DMS)]: the JT shield, inner vapor-cooled shield
(IVCS), middle vapor-cooled shield, and outer vapor-cooled shield (OVCS). The radiation
shields are cooled by the evaporated helium flowing through the vent-line plumbing.
Furthermore, the JT shield is cooled by a He-4 JT cooler, whereas the IVCS and the OVCS
are cooled by two double-stage Stirling coolers (2ST coolers). The first stage of the two
2ST coolers cools the OVCS, whereas the second stage cools the IVCS. These 2ST coolers are
shield coolers (SC). The other two 2ST coolers cool the inner DWR plumbing of the JT cooler.
These 2ST coolers are called pre-coolers (PC). In total, four 2ST coolers are used in the Resolve
cooling chain. The CMPs and cold heads (CH) of the 2ST coolers and the CMPs of the JT coolers
are mounted on the DMS.
Hitomi SXS development revealed that the vibration of the cryocoolers affects the detector’s
temperature stability and spectroscopic performance. The CHs and JT cooler produce only cyclic
vibration, so it was possible to avoid sensitive frequencies of the calorimeter by changing their
driving frequency. However, the CMPs produce not only cyclic vibration but also non-cyclic or
random vibration due to the ball bearings supporting the movement of the pistons. Therefore, a
Fig. 1 Cross-section of the Resolve Dewar.14
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Fig. 2 Cryogenic architecture of the Resolve Dewar.
VIS was added between each CMP and the DMS to mitigate this, decoupling the mechanical
vibration of the 2ST CMPs from the DMS.20 This improved the energy resolution. However, the
mechanical load applied to the 2ST CMPs during launch was only marginally within limits
because the VIS amplified the acceleration below 100 Hz.
For XRISM, it was decided to install VIS units for 2ST CMPs but not for JT CMPs and 2ST
CHs, as was done for Hitomi. The design of the VIS for Resolve incorporates a hold-and-release
mechanism (HRM) to prevent damage to the CMPs during launch or sine-vibe tests on the
ground. The VIS is an integrated system with a vibration isolator, thermal path, and
hold-and-release functions. The XRISM spacecraft was launched in September 2023, and the
functions of the VIS have been verified in orbit.
The 2ST and JT coolers installed in XRISM were developed by Sumitomo Heavy Industries
Ltd (SHI). To reduce wear on the lubrication material of the piston surfaces, ball bearings were
introduced as a guiding mechanism for the pistons. Although flexure springs were also considered for this purpose, ball bearings were selected for the compact design of the coolers. The first
2ST cooler featuring ball bearings was used in the cooling system of AKARI,4 which was
launched in 2006. At that time, vibration from the ball bearings was not identified as an issue.
Thus, based on the successful flight heritage of AKARI, the same 2ST design was adopted for
Hitomi. However, during the development of Hitomi, it became clear that vibration from the ball
bearings was problematic. Due to time constraints, a complete redesign of the 2ST was not feasible. Instead, countermeasures implemented by VIS proved successful, and the same approach
was applied to XRISM. The JT cooler with ball bearings was also utilized in JEM/SMILES,21
which was launched in 2009. However, due to issues related to impurity generation, a version
featuring flexure springs was developed and adopted for Hitomi. Currently, SHI is working on a
2ST cooler that incorporates flexure springs to minimize microvibration.22
2 VIS Requirements
The VISs are positioned between the DMS and four CMPs (SC-A, SC-B, PC-A, and PC-B). The
HRM, introduced for Resolve, has two states: released and locked. In the released state, the VIS
minimizes vibration transmission from the CMPs to the DWR, ensuring that the sensitive detector inside the DWR experiences negligible mechanical vibration. On the other hand, in the locked
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state of the HRM, the VIS can withstand the harsh mechanical environment during launch and
ground vibration testing without amplifying mechanical loads on the CMPs. In addition, the VIS
facilitates heat transfer from the CMPs to the DMS or a cooling plate, aiding in cooling the CMPs
during operation. In this section, essential requirements for the VIS are described.
2.1 Performance and Functional Requirements
2.1.1 Lifetime
As the XRISM satellite’s design lifetime is 3 years, the VIS must maintain its performance
requirements in orbit for at least 3 years. In addition, the VIS should withstand up to 3 years
of storage on the ground (under 1 G and in a normal atmosphere) before launch.
2.1.2 Vibration transmissibility
The primary characteristic of the VIS is its vibration transmissibility (τ), which is defined as the
ratio of the output force/moment (Fvo ) to the input force/moment (Fvi ). The definitions of these
forces and moments are shown in Fig. 3.
A lower transmissibility indicates superior vibration isolation. The requirement for vibration
transmissibility is the same as that for Hitomi. When the micro-disturbances due to the cryocoolers were studied for Hitomi, the vibration level threshold that produces the undesired temperature disturbance (thus causing detector performance degradation) was measured using the
Hitomi DWR engineering model in which the CSI engineering model was installed.23 Given
the similarity of the CSI and the DWR between Hitomi and XRISM, the same threshold level
was assumed for XRISM. When the threshold was determined, a six-degree of freedom (DOF)
vibrator was used to simulate the CMP vibration, and the allowable level within a specified
frequency range was determined.
The maximum permissible transmissibility was calculated as the quotient of the allowable
vibration level and the force acting on the CMP. Due to the performance characteristics of the
vibrators used, the defined frequency range was limited to 10 to 600 Hz for Fx and Fz and 100 to
600 Hz for the other axes. The requirement in the low frequencies of FY, MX, MY, and MZ,
which were not determined experimentally, was set by extrapolation. The transmissibility of the
VIS must not exceed the thresholds represented by the blue lines in Fig. 4.
We empirically knew that the detector was not significantly affected by vibrations below
100 Hz. As shown in Fig. 4, the required transmissibility for Fx and Fz in the frequency range
of 10 to 100 Hz was found to be ∼20 to 50. The passive isolator amplifies vibration most at its
resonant frequency, with the maximum value determined by the Q factor. As it was evident that
the resonant frequency of the isolator would be below 25 Hz and the Q factor would be below 10,
we assumed the extrapolated requirement is acceptable for frequencies below 100 Hz, other than
for Fx and Fz. Our assumption was finally verified by the verification test using the detector,
described in Sec. 4.6.2.
This requirement must be met both in orbit and in 1 G on the ground, even when the thermal
straps (TSs) and the cryocooler’s capillary tubes are attached. However, this requirement does not
apply in the locked state of the HRM.
CMP
z
x
y
CG of CMP
x
z
y
VIS
origin of Fvi
DWR
origin of Fvo
W=Fvo/Fvi
Fig. 3 Definition of vibration transmissibility (τ).
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Fig. 4 Requirement of VIS transmissibility.
2.1.3 Thermal conductivity
The VIS must be able to transfer a specified amount of heat from the CMPs for cooling. There are
two distinct thermal conductivity requirements that correspond to the different CMP heat transfer
paths. For SC-A, SC-B, and PC-A, with heat dissipated through the DWR, the thermal conductance between the CMP and the DMS should be at least 2 W∕K, with each CMP able to transfer
more than 63 W of heat. For PC-B, where the heat is dissipated to the cold plate cooled by a
spacecraft heat pipe, the thermal conductance between the CMP and the cold plate should be
equal to or exceed 1 W∕K.
2.1.4 Compressor load at vibration test
To prevent damage to a CMP mounted on the VIS, the maximum acceleration applied to the
CMP must not exceed the specified level for the CMP under acceptance test (AT) conditions.
These conditions include sine vibration, acoustics, shock, and quasi-static load. In other words,
the range of the acceleration amplification is specified. This requirement is related to the stiffness
requirements of the VIS in the locked state. If the stiffness is insufficient, the vibration transmitted to the CMP could be amplified, causing damage. This applies only when the HRM is
locked as the HRM is supposed to be locked whenever mechanical loads are applied.
2.1.5 Release operation requirements
The shock level generated by the HRM release must not exceed the CMP’s acceptance level. The
shock is also transmitted to the DWR, where the most sensitive component is a thin filter located
at the top of the DMS. Consequently, the allowable shock level at the top of the DWR is defined
as 100 Gsrs within the frequency range of 1000 to 4000 Hz.
During the release operation, the displacements of CMPs must remain within a specified
range. When the HRM is released, the CMP moves from the locked position to the released
position. Each CMP is connected to the corresponding CH via a capillary tube that facilitates
the flow of helium, the working fluid. In addition, SC-A and SC-B are equipped with loop heat
pipes (LHPs) to transfer heat to the spacecraft’s two radiator panels. Both the capillary tube and
the LHP are constructed from stainless steel and have maximum displacement requirements to
avoid plastic deformation. The allowable displacements are defined at the attachment points of
the capillary tube and the LHP, as shown in Table 1. This requirement must be met under both 1 G
conditions and in orbit.
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Table 1 Requirement on the maximum allowed CMP displacement during release.
Capillary attachment point
LHP attachment point
Item
Y (mm)
Z (mm)
Y (mm)
Z (mm)
SC-A
±7.8
±6.7
±5.0
±1.0
SC-B
±7.8
±6.7
±5.0
±1.0
PC-A
±1.9
±3.4
N/A
N/A
PC-B
±3.4
±8.3
N/A
N/A
2.2 Environmental Requirements
The VIS must withstand the mechanical loads during launch and ground testing (sine
vibration, acoustics, shock, quasi-static loads). The HRM was introduced to fulfill this
requirement. Fatigue-induced damage that can occur both on the ground and in orbit must
be considered.
Three sets of temperature ranges are defined for the VIS’s thermal environment requirements. The storage temperature range, within which no damage occurs, is defined as −60 to
þ 40°C. In addition, the performance assurance temperature range, within which the VIS operates properly, is defined as −30 to þ 40°C. Finally, the temperature range during launch and
ground testing is defined as þ 10 to þ 40°C.
The VIS must be compatible with the in-orbit environment in terms of vacuum, radiation,
atomic oxygen, and UV radiation.
The VIS must function under 1 G gravity (on the ground) and zero gravity (in orbit). In
particular, the vibration transmissibility requirement must be satisfied, and the pin-puller release
must operate without interference under both conditions.
2.3 Interface Requirements
The VIS must fit within the volume provided by the DWR design. Although this may seem
trivial, it imposes stringent constraints on the VIS design because the DWR design was established for the Hitomi SXS, and only minor modifications were possible.
The HRM pin-puller should be activated by Hitomi compatible electronics in orbit as no
changes to the electronics design are preferred. Therefore, the pin-puller should have a resistance
of 16.0 1.0 Ω (including the harness on the DWR) and must be released with a current pulse
ranging from 1.0 to 1.6 A for 30.0 1.0 ms. This is required for the in-orbit release operation of
the pin-puller as a spacecraft component applies the pulse. However, this requirement does not
apply when the HRM operates with ground support equipment (GSE).
2.4 Operational Requirements
On the ground, during the microcalorimeter performance test, the HRM must be released to meet
the transmissibility requirement. However, it must be locked during the vibration test to withstand acceleration. The HRM should be capable of locking and releasing at least 20 times during
ground testing without disassembling the satellite. Once the VISs are installed in the satellite,
they cannot be accessed manually; therefore, the HRM must be capable of being locked remotely
while on the ground. Meeting the thermal conductivity requirement during ground testing (except
during the thermal-vac test) is unnecessary as spot coolers can effectively cool CMPs under
atmospheric conditions.
In orbit, the VIS HRM should be released only once upon command from the ground. The
cryocoolers need to be activated as soon as possible after launch to preserve the cryogen (superfluid helium). Therefore, the HRM release operation should be executed immediately after
launch as it must occur before turning on the cryocoolers.
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CMP
CMP plate
TS-type C
Hold release mechanism (HRM)
(remotely resettable pin-puller)
TS-type B
TS-type C
DWR plate
Isolator strut (-Y)
Isolator strut (+Y)
TS-type B
HRM reset mechanism
Isolator strut (Z)
Fig. 5 VIS (SC-A, PC-A) exploded view.
3 Design of VIS
3.1 VIS Design Overview
Figure 5 is an exploded view of the VIS. The VIS comprises a DWR plate, a CMP plate, six
isolator struts, and TSs. The DWR plate incorporates an HRM release mechanism and a reset
mechanism. The CMPs are attached to this CMP plate. During ground performance tests or in
orbit after unlocking the HRM, the CMP is supported by the six isolator struts. However, the
HRM holds the CMP plate in place during launch or ground vibration tests. The HRM reset
mechanism, a unique feature of the XRISM VIS, remotely resets the previously released HRM.
Figure 6 shows the design variation of the VIS. The design of the VIS for SC-A and PC-A is
identical, whereas the design for SC-B is a mirror image of that for SC-A. For SC-A, SC-B, and
PC-A, four TSs connect the CMP and the DWR plate, whereas for PC-B, a single TS connects
the CMP to the cooling plate on the satellite.
3.2 Vibration Isolator Strut
Figure 7 shows the configuration of the vibration isolator struts. The CMP is supported over the
DWR plate by three pairs of vibration isolator struts (also referred to as isolator bipods). Figure 8
shows the structure of a single vibration isolator strut, comprising an upper component, a housing, two compression coil springs, and six-leaf springs made of single crystal shape memory
alloy (SCSMA).
Fig. 6 Three VIS designs: (a) for SC-A and PC-A, (b) for SC-B, and (c) for PC-B.
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Attachment to CMP
Flexible joint
Joint rod
Flexible joint
Attachment to DWR plate
Fig. 7 Configuration of vibration isolator struts.
Stage
Flexible joint
SCSMA leaf spring
SCSMA leaf spring
Joint rod
Compression coil spring
Center shaft
Upper component
Housing
Stage
Stage
Center shaft
Stage
Compression coil spring
Housing
Flexible joint
(a)
(b)
Fig. 8 Structure of vibration isolator strut: (a) perspective view, (b) conceptual diagram.
The compression coil springs and SCSMA leaf springs connect the upper component to the
housing. One end of the isolator strut is attached to the DWR plate using a flexible joint, whereas
the other end is connected to the CMP via a joint rod and a flexible joint. Initially, an attempt was
made to use ball joints to connect the struts; however, issues such as lubricant wear, particle
generation, and frictional instability were presented by the ball joints. Consequently, flexible
joints were opted for over ball joints.
Figure 9(a) shows the design of the flexible joint, which is made from A-286 stainless steel.
We modeled the flexible joint as a spring element using static analysis [Fig. 9(b)], and Table 2
shows the results. Using this modeled spring element, we conducted a modal analysis of the VIS
in the unlocked state (Fig. 10). In this model, the CMP was modeled as a mass. The results are
shown in Table 3. The label “pin joint” refers to using a three DOF constraint to connect the
struts, and “flexible joint” refers to using the spring element to connect them. As shown, the
Fig. 9 (a) Design of flexible joint. (b) Static analysis model of a flexible joint.
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Table 2 Stiffnesses of the designed flexible joint.
Lateral stiffness (N/mm)
100
Axial stiffness (N/mm)
352
Bending stiffness (Nm/deg)
1.2
Torsion stiffness (Nm/deg)
24.4
Z
X
Y
Fig. 10 Modal analysis model of VIS in an unlocked state.
Table 3 Result of modal analysis in unlocked state.
Frequency (Hz)
Mode
Pin joint
Flexible joint
Modal shape
1
4.3
3.5
Y rotation
2
6.8
6.6
X rotation
3
8.1
7.3
Z translation
4
10.2
9.6
X translation
5
11.5
10.3
Z rotation
6
11.6
11
Y translation
flexible joint lowers the modal frequencies compared with the ideal pin joint. This reduction is
due to the axial stiffness of the flexible joint; however, its effect was not critical.
The stiffness and natural length of the isolator strut can be adjusted by the two compression
coil springs. The natural frequencies of the suspended CMP were designed to be ∼4 to 12 Hz in
six degrees of freedom, thus avoiding the CMP’s drive frequency of 15 Hz. Furthermore, the
static displacements of the strut under 0 G and 1 G conditions are also important considerations.
Taking these conditions into account, the parameters of the compression springs were designed.
These struts have a simple lightweight structure, entirely made of metal materials without
using polymers such as rubber or oil. As a result, they exhibit extremely low outgassing
and aging.
3.2.1 SCSMA leaf spring
The SCSMA leaf springs act as damping elements.24–27 Shape-memory alloys (SMAs), widely
utilized in practical devices, exhibit remarkable properties such as shape-memory recovery and
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Stress, MPa
300
200
60 C
40 C
20 C
100
0
0
2
4
6
Elongation, %
8
10
Fig. 11 Stress–strain curve of SCSMA.28
4.0 mm
R= 9.5 mm
50.0 mm
T=0.8 mm
Fig. 12 Detailed design of SCSMA leaf spring.
superelasticity. Although typical SMAs are polycrystalline, the CuAl-based single-crystal shapememory alloy (SCSMA) has recently been developed.28 SCSMA possesses superior shapememory recovery and superelasticity. Superelasticity results from a stress-induced conversion
from austenite to martensite when stress exceeds a critical level and reversion from martensite
to austenite when stress falls below another critical level. These phenomena produce a pair of
constant stress plateaus in the stress–strain plot at specific temperatures. Single crystal superelasticity is characterized by an abrupt change in the slope of the stress–strain plot at a combination of the alloy’s stress, strain, and temperature characteristics. A typical stress–strain
property of SCSMA is shown in Fig. 11. For an elongation between roughly 2% and 8%, the
stress–strain curve exhibits hysteresis, which is used for damping.
Figure 12 is a detailed design of the SCSMA leaf spring. The thickness of the leaf spring (T )
and curvature radius (R) in the neutral position is 0.8 and 9.5 mm, respectively. Therefore, the
maximum strain of the leaf spring is T ∕2R ¼ 4.2% at which SCSMA shows significant
hysteresis.
3.2.2 Characteristics of isolator strut
A characteristic acquisition test was conducted on the breadboard model (BBM) of the isolator
strut. A displacement–force curve was obtained using a fatigue testing machine (Fig. 13). An
example of the displacement–force curve obtained in the test is shown in Fig. 14(a). The stiffness
(ks ) and the loss factor (ηs ¼ tan δ) are calculated as follows:
ks ¼
EQ-TARGET;temp:intralink-;e001;114;154
Fs1
;
Xs
ηs ¼ tan δ ¼
EQ-TARGET;temp:intralink-;e002;114;107
Fs0
:
Fs1
(1)
(2)
Here, Xs , Fs1 , and Fs0 represent the maximum displacement, the force at the maximum
displacement, and the force at zero displacement, respectively. The displacement dependency
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Fig. 13 Test configuration: (a) Isolator strut. (b) Fatigue testing machine. (c) Constant temperature
chamber.
Xs
(b) 20
4
N/mm
0
10
0.1
5
–4
–8
–1.5
0
–1
–0.5
0
0.5
1
0
0
1.5
0.2
Displacement (mm)
1 mm
0.5 mm
0.1 mm
0.05 mm
0.6
0.8
1
Displacement (mm)
0.23 mm
(c) 20
stiffness (w/o spring)
stiffness (w/ spring)
loss factor (w/o spring)
loss factor (w/ spring)
(d) 20
0.2
0.2
15
0.1
N/mm
10
η (tanδ)
15
N/mm
0.4
10
0.1
η (tanδ)
Force (N)
15
Fs1
Fs0
0.2
η (tanδ)
(a) 8
5
5
0
0
–60
0
0
5
10
frequency (Hz)
stiffness (w/o spring)
loss factor (w/o spring)
15
20
0
–40
–20
0
20
40
60
80
temperature (°C)
stiffness (w/ spring)
loss factor (w/ spring)
stiffness
loss factor
Fig. 14 (a) Displacement-load plot of single isolator strut. (b) Displacement dependency (15 Hz,
20°C). (c) Frequency dependency (0.23 mm, 20°C). (d) Temperature dependency (without spring,
15 Hz, 0.23 mm).
is shown in Fig. 14(b). The test was conducted with two configurations: one with the coil springs
and one without them. As the displacement increases, the stiffness decreases, and the loss factor
increases. As the coil springs have no damping property, the strut with the coil springs has a lower
loss factor than the strut without the coil springs. The frequency dependency, shown in Fig. 14(c),
is relatively small, within the 5- to 20-Hz range. The temperature dependency, obtained only for
the strut without coil springs, is shown in Fig. 14(d). As the temperature increases, the stiffness
increases, and the loss factor decreases. At temperatures below −30°C, the SCSMA is in the
temperature-induced martensitic state.
The isolator strut exhibits damping properties at low temperatures, such as −54°C. The availability of damping materials that can be used at the same temperatures as this strut is limited.
These data are used to correlate the structure math model, confirming that the VIS design should
satisfy the transmissibility requirements shown in Fig. 4.
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An accelerated lifetime test was conducted on the SCSMA leaf spring. The spring was subjected to deformation at 200 Hz, which is 13.3 times the expected frequency in orbit. After three
months of testing, no significant changes were observed in its properties, indicating a lifetime of
more than 3 years in orbit. This confirms that the requirements stated in Sec. 2.1.1 have been met.
3.3 Hold and Release Mechanism
The HRM holds the CMP securely during launch to prevent excessive movement and avoid VIS
malfunction. The HRM for the XRISM VIS is integrated into the DWR plate and is in either the
locked state or the released state, as depicted in Fig. 15. In the locked state [Fig. 15(b)], the CMP
and DWR plates are rigidly connected, allowing the VIS to withstand high acceleration. The
inner conical fitting attached to the DWR plate presses against the outer conical fitting at the
four corners of the CMP plate, securing it in place. The preload, which pushes the inner conical
fitting against the outer conical fitting, is a critical parameter in the design of the HRM. The VIS
design should ensure the following:
•
During launch or ground tests, the preload must exceed the maximum force applied to the
engaged fittings. Otherwise, a gap may form between the outer and inner conical fittings
during vibration, which increases CMP acceleration.
• The structure of the HRM, including the CMP plate and the DWR plate, should be stiff
enough to prevent gaps in the fittings under the applied load and strong enough to avoid
permanent deformation.
• The HRM reset mechanism should generate sufficient force to apply the preload.
In addition, consideration should be given to the material used in the HRM, the size of the
motor and pin-puller, and the overall mechanism design to meet these conditions.
In the released state [Fig. 15(c)], three isolator bipods suspend the CMP from the DWR
plate, reducing vibration transmission from the CMP to the DMS.
The transition from the locked state to the released state is triggered by activating the EBAD
TiNi™ auto-resettable P5-5004 pin-puller, which can be remotely reset. When the pin-puller is
activated, the drive shaft moves to the released position due to spring-restoring force, separating
the outer and inner conical fittings, and putting the VIS in the released state.
Conversely, the HRM reset mechanism returns the VIS to the locked state. Its motor moves
the drive shaft from the released to the locked position. Subsequently, the pin-puller operates to
(a)
Auto-resettable pin-puller
2ST compressor (CMP)
Geared motor
CMP plate
CMP fixes to the CMP plate
DWR plate is
fixed to the DWR
Isolator bipod
CMP is supported by 3 bipods
in released state
(b)
CMP side
Conical-conical
fitting
DWR side
Motor arm
Drive shaft
(c)
Pin-puller
rod
Compression
spring
Fig. 15 HRM: (a) Schematic view. (b) Locked state. (c) Released state.18
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lock the mechanism, and the pin-puller rod fits into the drive shaft groove, securing the VIS.
Finally, the motor arm moves to the right side to avoid interference when the VIS is released.
GSE controls the HRM reset mechanism, as there is no need for in-orbit operation. Note that the
simple release operation by a pin-puller and springs provides robust release in orbit, in contrast to
the more sophisticated on-ground reset operation utilizing a motor and gears. The HRM reset
mechanism is designed to lock and release multiple times on the ground, facilitating Resolve’s
performance evaluation. The HRM was designed so that the mechanism can be held in place by
the motor alone, without resetting the pin-puller, whereas the pin-puller can be locked as necessary. Both operations can be performed electrically without human access. In addition, two heaters (primary and redundant) were attached to a pin-puller for temperature control as the
requirement in Sec. 2.2 specifies operation in the range of −30°C to þ 40°C.
3.4 Thermal Strap
3.4.1 Function and design of TS
The TS is a heat path used to dissipate the heat generated by the CMP. It is designed to be flexible
to prevent the transmission of CMP vibrations and is composed of multi-layer graphite sheets
(GS) and aluminum end fittings. The GS material used is called Graphinity™, which is manufactured by Kaneka Corporation. Graphinity™ is known for its lightweight (with a density of
1.8 g∕cm3 ) and high thermal conductivity (three times that of copper). In addition, it is soft, easy
to bend, and has low outgassing properties.
Three types of TSs, namely A, B, and C, have been developed for the XRISM VIS, as shown
in Fig. 16. Thermal conductivity and stiffness tests were performed on these types. Type A was
used in conjunction with PC-B, which connects the CMP and the satellite base’s cold plate.
Types B and C, on the other hand, were used with SC-A, SC-B, and PC-A, which connect the
CMP and the DMS.
3.4.2 Verification of TS
The thermal conductance of the TSs was measured in the configuration shown in Fig. 17. One
end of the TS, which is expected to be hotter in orbit, was attached to a heater, whereas the other
Fig. 16 Three types of TS: (a) Type A. (b) Type B. (c) Type C.
Heater
TS
Cold plate
Vacuum chamber
(a)
(b)
Fig. 17 Thermal conductivity test configuration of TS: (a) Measurement setup. (b) Vacuum
chamber.
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Table 4 Results of thermal conductance tests of individual TS.
ΔT (K)
Thermal
conductance
(W/K)
Requirement
(W/K)
Assigned to
Type
Phase
No.
Heater
power (W)
A
FM
1
29.7
21.6
1.37
1.00
PC-B
B
QM
1
29.7
35.8
0.83
0.75
—
QM
2
29.7
30.1
0.99
—
FM
1
29.7
33.7
0.88
SC-A
FM
2
29.7
37.3
0.80
SC-B
FM
3
29.7
30.9
0.96
PC-A
FM
4
29.7
32.4
0.92
SC-B
FM
5
29.7
30.4
0.98
PC-A
FM
6
29.7
32.9
0.90
SC-A
QM
1
22.4
32.0
0.70
QM
2
22.4
36.4
0.61
—
FM
1
22.4
34.4
0.65
SC-B
FM
2
22.4
33.0
0.68
SC-B
FM
3
22.4
32.3
0.69
PC-A
FM
4
22.4
31.6
0.71
PC-A
FM
5
22.4
34.7
0.64
SC-A
FM
6
22.4
35.0
0.64
SC-A
C
—
0.62
end was attached to a cold plate. The measurement was conducted in a vacuum chamber, and the
thermal conductance was calculated as the ratio of the heater power to the temperature difference
between both ends. All TSs were found to meet the requirement stated in Sec. 2.1.3. Table 4
shows the results of the thermal conductance tests of TSs. As PC-A was the most thematically
critical, the most conductive straps were assigned to it. The remaining straps were assigned to
SC-A and SC-B to ensure that each system’s thermal conductance became nearly equal.
The stiffness of the TSs was measured using the configuration shown in Fig. 18. A vibrator
was attached to one end of the TS, and the displacement was measured with a laser displacement
sensor. The reaction force at the opposite end of the TS was also recorded. The stiffness of the
TSs was determined as the ratio of the measured force to the displacement. It was confirmed that
the stiffness of the TSs was significantly lower than that of the isolator struts, ensuring that the
Laser displacement sensor
(a)
Force sensor
(b)
TS
Vibrator
(c)
Fig. 18 Stiffness test configuration of TS: (a) Measurement in the longitudinal direction.
(b) Measurement in the traverse direction. (c) Measurement at low temperature (−30°C).
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vibration isolation performance was not impeded. The stiffness measurement test for the BBM
was conducted at low temperature. The TS was cooled using a freeze spray, as shown in
Fig. 18(c), and the stiffness was measured when it reached −30°C. The stiffness was found
to be 1.3 times higher at −30°C compared with 20°C (room temperature). However, this increase
is not considered significant in terms of vibration isolation performance. In addition, stiffness
was found to exhibit minimal frequency dependence.
4 Verification of VIS
4.1 Overview
During the manufacturing phase, five VISs were produced, including one qualification model
(QM) and four flight models (FMs): SC-A, SC-B, PC-A, and PC-B. The following verification
tests were conducted for all models. Although vibration transmissibility, thermal conductance,
and mechanical endurance are not significantly affected by gravity, the displacement of the
released HRM is greatly influenced by it. Therefore, the HRM release operation under 1G was
verified through the HRM release operation test, and the operation under 0-G was verified by a
combination of this test and analysis.
4.2 Vibration Transmissibility
4.2.1 Method of measurement
The requirement for vibration transmissibility is defined as the ratio of the input force/moment at
the CMP to the reaction force/moment of the fixed end, as described in Sec. 2.1.2. However,
fixing one end of the VIS in inertial space is quite challenging, especially when measurements at
various temperatures and directions of gravity is necessary. Therefore, the transmissibility was
calculated as the ratio of the base acceleration/angular acceleration to the CMP acceleration/
angular acceleration. The base fixture was excited using an inertial vibrator, and the acceleration
of a CMP simulator was measured, as shown in Fig. 19. Two inertial vibrators were used for each
measurement, as indicated by the blue arrows; the red arrows show the movements of the CMP.
For the translation measurement, two vibrators were driven in the same direction, whereas for the
rotation measurement, they were driven in opposite directions.
The following explanation outlines how the same transmissibility can be obtained by measuring the ratio of acceleration to acceleration instead of the ratio of force to force. Let us consider
a single-degree-of-freedom (1-DOF) vibration isolator model, as depicted in Fig. 20.
Figure 20(a) shows a fixed base model where the excitation force acting on the mass is a
sinusoidal force (Fmi expðjωtÞÞ with angular frequency ω. The equation of motion and the reaction force at the base (Fmo expðjωtÞ) are as follows:
X translation
Y translation
Z translation
Y rotation
Z rotation
Z
X
Y
VIS local coordinate
X rotation
Fig. 19 Transmissibility test configurations.
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(a)
exp
exp
(b)
mm
km
exp
mm
cm
cm
km
exp
exp
Fig. 20 1-DOF model of vibration isolator: (a) fixed base configuration and (b) base excitation
configuration.
ð−mm ω2 þ jcm ω þ km ÞX a expðjωtÞ ¼ Fmi expðjωtÞ;
(3)
Fmo expðjωtÞ ¼ ðjcm ω þ km ÞX a expðjωtÞ:
(4)
EQ-TARGET;temp:intralink-;e003;114;594
EQ-TARGET;temp:intralink-;e004;114;564
Here, mm , cm , km , and X a represent mass, damping coefficient, spring constant, and displacement amplitude of the mass, respectively. The vibration transmissibility τm, which is the
ratio Fmo to Fmi , becomes as follows:
τm ¼
EQ-TARGET;temp:intralink-;e005;114;521
Fmo
jcm ω þ km
¼
:
Fmi −mm ω2 þ jcm ω þ km
(5)
On the other hand, for the base excitation model shown in Fig. 20(b), the equation of motion
is given by
−mm ω2 Xb ejωt þ jcm ωðX b − Y b Þejωt þ km ðXb − Y b Þejωt ¼ 0:
EQ-TARGET;temp:intralink-;e006;114;460
(6)
Here, X b and Y b represent the displacement amplitude of the mass and the base, respectively.
From Eqs. (5) and (6), the ratio of the mass acceleration (ω2 X b ) to the base acceleration (ω2 Y b ) is
as follows:
EQ-TARGET;temp:intralink-;e007;114;399
ω2 X b X b
jcm ω þ km
¼
¼
¼ τm :
ω2 Y b Y b −mm ω2 þ jcm ω þ km
(7)
As explained above, the two test methods are mathematically equivalent. However, there are
a few points that should be mentioned:
•
The base fixture should be sufficiently rigid to prevent resonance within the measurement
frequency range.
• In general, vibrators and accelerometers perform poorly at low frequencies. It should be
noted that measurements at low frequencies may be challenging.
• The VIS might have displacement dependency. Therefore, to compare vibration transmissibility measured by different methods, it should be measured at the same relative displacement in the same direction.
4.2.2 Measurement result
Vibration transmissibility tests were conducted on five VISs at three temperatures: room temperature (RT), low temperature (−30°C), and high temperature ( þ 40°C). Due to the characteristics
of the vibrator and the accelerometers, the measurement frequency range was limited to 25 to
600 Hz. The base fixture was designed to be sufficiently rigid to avoid interfering with the test
results.
The vibration transmissibility requirement must be met in orbit and under 1 G on the ground.
However, the tests were conducted exclusively under 1 G conditions because testing under 0 G
conditions is challenging in a ground-based environment, and the vibration transmissibility is not
to be affected by gravity as long as the springs used in the vibration isolator have linear characteristics. Specifically, the SC-A and SC-B VIS units are mounted on the side of the DMS, with
gravity acting in the Y direction of the local coordinate system. Conversely, the PC-A and PC-B
units are mounted on the bottom of the DMS, experiencing gravity in the −Z direction. These
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Fig. 21 Transmissibility test configurations: (a) SC-A and SC-B. (b) PC-A. (c) PC-B.
Fig. 22 Vibration transmissibility of the qualification model (QM).
configurations are shown in Fig. 21. For the QM, the test was performed with gravity acting in
the þ Z direction (not shown in Fig. 21). Notably, the transmissibility did not exhibit significant
differences based on the direction of gravity.
Figure 22 illustrates the transmissibility of the QM across all temperatures, confirming that
the requirement stated in Sec. 2.1.2 was met. All FMs also met the requirement.
4.3 Thermal vacuum test
The thermal conductance of each VIS was verified with a full set of TSs. The test was conducted
in a vacuum chamber under the expected heat load in orbit. Heat generation from the CMP was
simulated using a heater attached to the CMP simulator. All VISs met the requirement specified
in Sec. 2.1.3, as shown in Table 5.
Thermal cycle tests were conducted in a vacuum chamber for all the FMs and the QM. First,
the VISs were exposed to the lowest and highest storage temperatures (−60 to þ 40°C). Second,
an eight-round thermal cycle test was conducted between the lowest and highest operating temperatures (−30 to þ 40°C). After these tests, all the VISs operated normally and met the requirements stated in Sec. 2.2.
4.4 HRM Function
The operation of the pin-pullers’ release was verified in a vacuum chamber. The tests were conducted under worst-case conditions, including a temperature of −30°C and a current of 1.0 A.
The pin-puller operates using the shape memory effect of the SMA, which is triggered by a rise in
temperature when an electrical current is applied. Therefore, operation at the minimum temperature and current is considered the most challenging. All of the pin-pullers met the requirements
in Sec. 2.3, as shown in Table 6.
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Table 5 Results of thermal conductance test.
Item
Requirement
(>W/K)
Thermal
conductance (W/K)
Heat load (W)
High temp. (°C)
Low temp. (°C)
SC-A
2.0
2.42
63.2
40.3
14.2
2.44
34.9
39.5
25.2
2.46
63.6
42.9
17.0
2.49
34.6
40.6
26.7
2.73
63.3
37.2
14.0
2.78
35.0
37.7
25.1
1.04
64.4
37.0
-25.1
1.05
34.9
38.6
5.3
2.14
63.7
41.4
11.6
2.14
34.6
42.0
25.9
SC-B
PC-A
PC-B
QM
2.0
2.0
1.0
2.0
Table 6 Results of the pin-puller operation test under a temperature of −30°C.
Current (A)
Duration (ms)
Item
Result
Requirement
Result
Requirement
Result
SC-A
1.0
1.0 to 1.6
14.0
<29
Pass
SC-B
1.0
18.4
Pass
PC-A
1.0
16.4
Pass
PC-B
1.0
17.2
Pass
QM
1.0
19.2
Pass
displacement (mm)
The displacements of the CMP were verified. The displacements under 1G were measured
using laser displacement sensors. When the lock was released, the CMP was pulled by gravity,
causing it to oscillate with damped motion. An example of the displacement time history is
shown in Fig. 23, demonstrating that the movement of the CMP stabilizes in ∼1 s.
The displacements under 0G were verified based on the 1G test and the 1G static analysis of
the VIS. The static analysis indicates a displacement of 0 mm at 0G, whereas the displacement of
the FMs at 0G is adjusted to minimize the displacement at 1G. Therefore, it can be estimated that
the difference between the 1G static analysis and the settled displacement at 1 G represents the
2
1.6
1.2
0.8
0.4
0
0
1
2
3
time (s)
4
5
6
Fig. 23 SC-A CMP displacement in the Y direction at the capillary mounting point.
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Table 7 CMP displacement during the releasing operation.
Item
Measured
point
SC-A
Capillary attachment point
LHP attachment point
SC-B
Capillary attachment point
LHP attachment point
PC-A
PC-B
Capillary attachment point
Capillary attachment point
Analysis
1G Disp. 1G Disp. 0 G Disp.
(1G static) Requirement
(max)
(steady) (estimated)
Direction
(mm)
(mm)
(mm)
(mm)
(mm)
Y
2.19
±7.8
1.81
1.12
−1.07
Z
0.38
±6.7
0.41
0.25
−0.13
Y
2.74
±5.0
2.27
0.73
−2.01
Z
0.61
±1.0
0.66
0.41
−0.20
Y
2.19
±7.8
1.76
1.11
−1.08
Z
0.38
±6.7
0.45
0.27
−0.11
Y
2.74
±5.0
2.20
0.79
−1.95
Z
0.61
±1.0
0.73
0.44
−0.17
Y
−0.07
±1.9
−1.07
−0.64
0.43
Z
0.77
±3.4
1.15
0.72
0.05
Y
−1.07
±3.4
−0.92
−0.31
0.76
Z
0.77
±8.3
1.62
1.07
0.30
displacement of the FM at 0G. All results are summarized in Table 7, where “1G Disp. (max)”
refers to the peak displacement, “1G Disp. (steady)” indicates the settled displacement, and “0G
Disp. (estimated)” represents the estimated displacement under 0G. As shown, all displacements
comply with the requirements outlined in Sec. 2.1.5.
For QM, a lifetime test of the HRM mechanism was conducted. As described in Sec. 2.4, the
HRM must be capable of being locked and released at least 20 times during ground testing without disassembling the satellite. During QM testing, after the locking and release process was
performed at least 40 times (twice the requirement), it was confirmed that the mechanism was
still in good condition.
4.5 Vibration Test
Vibration tests were conducted to verify the mechanical endurance of the VIS during launch,
especially when the HRM is locked. These tests also confirmed that the acceleration at the
CMP did not exceed its acceptable level during the AT. The tests were performed using a
CMP simulator that emulated the mass properties of the actual CMP.
Both sine and random vibration tests were conducted in the X, Y , and Z directions. Modal
tests were carried out before and after each vibration test to ensure that no damage occurred.
Figure 24 shows the frequency response function (FRF) of the QM VIS during the modal test.
The FRF, which is the ratio of the acceleration measured at the CMP simulator to the base, shows
that there was no significant amplification under 100 Hz. Table 8 and Fig. 25 show the results of
Fig. 24 Frequency response function of VIS QM model at the CMP simulator.
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Table 8 Modal analysis in a locked state.
Effective mass ratio (%)
Mode
Hz
X
Y
Z
Rx
Ry
Rz
1
111.7
0
26
0
91.5
0
6.2
2
141.0
40.1
0
0
0
63.3
0
3
159.3
0
0
81.1
0
10.1
0
4
195.5
39.8
0
0
0
0.3
0
5
254.5
0
54.8
0
1.3
0
13
6
474.1
0
0
0
0
2.2
0
7
586.6
0
0
0
0
0
20.6
8
587.6
0
0
4.5
0
5.1
0.1
9
738.2
0.5
0
1.4
0
2.5
0
10
742.0
0.1
0
1.2
0
0.1
0
11
766.2
0.4
0
2.7
0
0.1
0
12
893.9
0
0
1.2
0
2.7
0
81.5
90.1
95
95.8
88
50
Total (%)
(a)
(b)
Mode 1 111.7 Hz
(c)
Mode 2 141.0 Hz
Mode 3 159.3 Hz
Fig. 25 Modal shapes in launch configuration. (a) Mode 1 111.7 Hz. (b) Mode 2 141.0 Hz. (c) Mode
3 159.3 Hz.
the modal analysis in the locked state and the modal shapes of the first three modes, respectively.
The CMP was modeled as a single mass. The peaks in Fig. 24 are likely to correspond to
these modes.
Sine vibration tests were performed on the FMs at the AT level, whereas the QM was subjected to vibration at the QT level, which is 1.25 times as high as the AT level. The random
vibration test level was the same for both QM and FM, but the duration was twice as long for
QM than for FM. For the QM, one article was tested at a time (Fig. 26). However, to expedite the
test campaign, two FMs were tested simultaneously (Fig. 27). All the articles met the requirements outlined in Secs. 2.1.4 and 2.2.
4.6 Verification and Validation at the Instrument and Spacecraft Levels
The four sets of VIS were integrated into the DWR, and verification was conducted at the instrument level. Following the integration of the Resolve instrument, a series of spacecraft-level verification tests, including electrical, thermal, and mechanical tests, were carried out. These
verifications at both the instrument and spacecraft levels confirmed the performance and
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Fig. 26 Vibration test configuration for QM VIS.
Fig. 27 Vibration test configuration for FM VIS.
functionality of the VIS. The system’s vibration isolation efficacy was substantiated by measurements of the vibration levels on the DMS and analysis of the detector noise spectra and energy
histograms derived from calibration with X-ray sources.
In addition, the release and locking functions of the HRM were tested repeatedly. The thermal performance of the TS was demonstrated during the spacecraft thermal-vac test.
4.6.1 Vibration level on the DMS
To verify the effect of the VIS before installing the detector in the DWR, the allowable vibration
level at the representative point of the DMS was defined as a requirement based on the Hitomi
results.18 This requirement applies not only to the VIS but also to the entire DWR system, which
includes the CMPs, CHs, and VISs. The allowable vibration level in the Z direction (DWR local
coordinate system shown in Fig. 1) at the top of the middle of the DMS (between SC-A and SCpffiffiffiffiffiffi
B) was defined as 2 × 10−4 G∕ Hz or lower within a frequency range of 100 to 600 Hz. This
requirement was applied to the continuum component and not to the line components at the
harmonics of the cryocooler’s drive frequency. As shown in Fig. 28, the continuum component
pffiffiffiffiffiffi
of the vibration level when the cryocoolers were operated was below 1 × 10−4 G∕ Hz, which
satisfies the requirement. Both CMPs and CHs produce many harmonics of their driving frequency of 15 Hz. However, the harmonics shown in Fig. 28 are mostly caused by CHs.
Although CMPs are mounted on the VIS, CHs are rigidly mounted on the DWR. Note that these
harmonics do not degrade the detector performance as long as the frequencies are off the natural
frequency of the CSI or the detector.
4.6.2 Detector performance
The four VIS units were installed on the DWR before the detector performance test. The performance of the VIS was demonstrated by measuring the detector noise and energy resolution. In
all the detector performance tests on the ground, regardless of ambient or thermal-vac environment (i.e., during the spacecraft thermal-vac test), the noise feature was free from vibration disturbance when the VIS was released, validating the performance of the VIS. The energy
resolution of the detector was <5 eV at 5.9 keV, satisfying the ground test allocation
(<5.9 eV) for the in-orbit requirement of <7 eV. The detector noise and energy resolution
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Fig. 28 Power spectral density (PSD) of the acceleration measured on the DMS when all the coolers are operating.18
performance were also measured with the VIS in the locked state, where vibration isolation was
not applicable. The noise was excessive at low frequencies, and the energy resolution was
degraded. Examples of the noise spectra and energy histogram of the Fe-55 calibration source
detected by detector pixels are shown in Fig. 29. Figures 29(a) and 29(b) show the data taken with
the VIS released, whereas Figs. 29(c) and 29(d) show those with the VIS locked. The noise spectra
represent the Fourier-transformed detector voltage signal when no X-ray hits occur. When the VIS
was locked and the cryocooler vibration was not isolated, the noise spectrum exhibited excess in
the low-frequency band, affecting the detector’s energy resolution. The low-energy excess suggests
that the detector and/or its wiring stochastically move, leading to variations in detector gain. The
two peaks of the energy histogram in panel (b) correspond to the Ka1 and Ka2 transitions of Mn
Fig. 29 Comparison of noise spectra (a), (c) and energy histogram (b), (d). Panels (a) and (b) are
taken with the VIS released, whereas panels (c) and (d) are taken with the VIS locked. Blue, red,
green, and magenta lines in panels (a) and (c) show the noise spectrum of pixels 0, 1, 2, and 3,
respectively. Energy histograms (b), (d) are taken with the Fe-55 calibration source. Data from all
detector pixels are combined.
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Table 9 Number of operations of the HRM.
Environment
On-ground (instrument-level
and spacecraft-level)
On-orbit
HRM state
SC-A
SC-B
PC-A
PC-B
Pin-puller release
23
24
30
26
Lock (motor only)
7
7
8
7
Lock (pin-puller)
23
24
30
26
Pin-puller release
1
1
1
1
Lock (motor only)
N/A
N/A
N/A
N/A
Lock (pin-puller)
N/A
N/A
N/A
N/A
atoms produced by electron capture of the Fe-55 radioisotope. The distinction between the two
peaks is unclear in panel (d) due to the degradation of energy resolution from ∼4.0 to ∼6.5 eV.
4.6.3 HRM function
The HRM was repeatedly released and locked. When testing the detector’s performance, the
HRM was in the released state. By contrast, the HRM had to be locked during mechanical tests
of the DWR and spacecraft. In addition, as the CMP shifts ∼2 mm under 1 G, the HRM should be
locked to prevent excessive movement when relocating the DWR or the spacecraft. Once the
DWR was integrated into the spacecraft, the VIS became inaccessible to the operational staff.
The HRM was designed to be electrically reset to the locked position, obviating the need for
manual intervention.
Table 9 summarizes the HRM operations, underscoring the significance of the reset function
without needing to replace the pin-puller unit. Without this function, the financial and scheduling
repercussions would be considerable, possibly impeding the comprehensive validation of detector performance. The HRM can secure the CMP solely with its motor mechanism, bypassing the
need to reset the pin-puller. The pin-puller was reset for configurations akin to actual flight conditions (e.g., mechanical tests, thermal-vac tests, and HRM functionality tests), whereas the
motor-lock setup was employed to secure the mechanism during spacecraft transport operations.
The reset process was intended to be executed using specialized GSE. The XRISM VIS cannot
reset the HRM while in orbit.
The spacecraft’s thermal-vac test offered the optimal opportunity to demonstrate the HRM’s
release feature in an in-orbit-like environment. As the pin-puller employs an SMA, its energy
margin decreases at lower temperatures. As a demonstration in the worst case, the release procedure was conducted at a low spacecraft temperature, with the heater regulated to align with inorbit operation protocols. The successful release of the four HRM units validated their functionality and their compatibility with the spacecraft’s drive electronics.
4.6.4 Thermal strap performance
The thermal performance of the VIS, achieved by the TS units, was demonstrated during the
spacecraft thermal-vac test. However, due to the limited number of temperature sensors, accurate
measurements of each TS were not possible. Table 10 shows the temperatures of the four CMPs,
the allowable temperature ranges, and the expected values from the math-model-based analysis.
The data shown correspond to the hot case with nominal cryocooler power. The CMP temperatures were within the acceptable range with a fair margin, thereby demonstrating the performance of the TSs.
4.6.5 On-orbit performance
The HRM of the VIS was successfully released on September 7, 2023, at 0:33 UT, ∼50 min after
launch, with the temperature of the pin-pullers controlled at 30°C. The Resolve detector was
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Table 10 Temperature measured during the spacecraft thermal-vac test (hot case).
Position
Measured temperature (°C)
Analysis prediction (°C)
Allowable range (°C)
SC-A CMP
+0.4
−3.0
−40/+40
SC-B CMP
+1.6
−3.0
−40/+40
PC-A CMP
+27.8
+23.0
−40/+40
PC-B CMP
+14.7
+9.0
−40/+40
Fig. 30 (a) An in-orbit noise spectrum of the Resolve detector. (b) Energy histogram.29
cooled to its operating temperature of 50 mK on October 9, 2023. The detector’s performance
was then verified during the commissioning period of the mission.
The detector’s energy resolution met the requirement of 7 eV, achieving ∼5 eV. The temperatures of the CMPs are all within the allowable range by a fair margin. The typical temperatures
of each CMP measured in orbit are −10°C for SC-A and SC-B, þ 10°C for PC-B, and þ 25°C
for PC-A, with a variation of ± a few °C depending on the orbital motion and pointing direction.
The temperature of the VIS should be similar to that of the associated CMP. The VIS functions in
all required aspects. Figure 30 shows an example of the noise spectra and the detector energy
histogram taken in orbit.
5 Lessons Learned
The following lessons were learned through the development of the XRISM VIS and are helpful
when developing a VIS to decouple microvibrations from cryocoolers with a sensitive detector.
•
In general, highly sensitive sensors are also susceptible to noise sources such as mechanical
vibrations, a broad band of electromagnetic waves, and others. The most reliable way to
determine noise immunity is to test using real (similar) hardware. Therefore, we strongly
recommend developing an engineering model (EM) of the sensor in the early stage of
development. In the case of Hitomi, the mechanical noise issues were not recognized until
testing was conducted using the EM. For XRISM, as the designs of the CSI and the DWR
were quite similar, the same threshold level could be assumed.
• In the case of Hitomi, the decision to adopt the VIS was made during the manufacturing
phase. As a result, the interface between the VIS and the DWR had to be aligned with the
already manufactured DWR, imposing rigid constraints on the design of the VIS. In contrast, for Resolve, although the fundamental DWR design is the same as that for the SXS,
the decision to adopt VIS was made before the manufacturing of DWR, allowing for the
redesign of the DWR to create an interface tailored to the VIS. This significantly increased
the design flexibility of VIS and provided a slight buffer in the development schedule. If the
DWR could have been newly designed, the VIS would have been designed much more
flexibly. If mechanical vibrations could cause problems, the need for VIS should be
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determined early in the project. If uncertainty remains, the development should proceed
with the incorporation of the VIS as it can be easily removed afterward.
• The detector’s performance in the in-orbit configuration can only be measured when the
HRM (i.e., the launch-lock mechanism) is released. However, it must be locked for
mechanical verification or other purposes (e.g., for protection from damage due to large
movement under 1 G). As these two requirements conflict, the capability to re-lock the
mechanism without replacing any parts is necessary. The design described in this paper
provides one such solution.
• As the pin-puller is based on an SMA, the energy required to release it depends strongly on
the temperature of the pin-puller housing. To reduce uncertainty and secure the energy
margin, the thermal design of pin-pullers, which includes thermal isolation and heater control, is crucial.
• An HRM reset failure occurred during one of the spacecraft-level tests. Although it was
impossible to directly observe the VIS, the cause of the failure was determined to be a slight
positional deviation of the HRM’s parts as identified through a fault tree analysis. The
following measures should be taken to avoid such failures: (1) The tolerances for moving
parts should be large enough to accommodate slight deviations in their relative positions.
(2) For components that bear a large load in the shear direction of the fastening bolts,
positioning should be done using shear pins in conjunction with the bolts.
6 Conclusion
The design and performance of the VIS for the Resolve instrument onboard the XRISM spacecraft are described. The VIS was designed to mitigate the degradation of the spectroscopic performance of a cryogenic microcalorimeter instrument due to mechanical disturbances from
cryocooler compressors. The VIS includes a vibration isolation function, a heat transfer function,
and hold-and-release functions. These functions work together to keep the cryocoolers within
their allowable environment and ensure that there is no negative impact on the detector’s performance. We hope that the design solutions and lessons described in this paper can serve as a
valuable reference for future cryogenic space missions, particularly those requiring unprecedented high sensitivity.
Disclosures
The authors declare the following financial interests and/or personal relationships that could be
considered potential competing interests: Kazuki Watanabe holds patents related to a shockabsorbing device issued to WEL Research. All other authors declare that they have no known
competing financial interests or personal relationships that could have influenced the work
reported in this paper.
Code and Data Availability
The data supporting the findings of this article originate from internal reports generated during the
project and are not publicly available. Code sharing is not applicable to this article as no new code
was developed.
Acknowledgments
The authors express their gratitude to all members of the Hitomi SXS team and the XRISM Resolve
team. During the preparation of this work, the authors utilized Microsoft Copilot and OpenAI
ChatGPT to enhance the readability and language of the paper. After using these services, the
authors thoroughly reviewed and edited the content as needed and take full responsibility for the
published article’s content.
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Susumu Yasuda is a manager for structural engineering at the Japan Aerospace
Exploration Agency (JAXA). He received his BS and MS degrees in mechanical engineering
from the University of Tokyo in 1991 and 1993, respectively. From 1993 to 2006, he worked
at Canon Inc., and joined JAXA in 2006, where he contributed to the development of vibration
isolators for Hitomi/SXS and XRISM/Resolve. His primary research interest lies in mechanical
vibration in spacecraft.
Yoh Takei is an instrument engineer at the Japan Aerospace Exploration Agency (JAXA). He
received his BS, MS, and PhD degrees in science from the University of Tokyo in 2001, 2003,
and 2006, respectively. With over 15 years of experience, he has developed cryogenic X-ray
spectrometers for the SXS instrument on Hitomi and the Resolve instrument on XRISM. In
addition to instrument engineering, he has also worked as an astrophysicist and spacecraft systems engineer.
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Atsushi Okamoto is a manager in the thermal control technology division at JAXA. He received
his BS, MS, and PhD degrees in aerospace engineering from Kyushu University in 2001, 2003,
and 2018, respectively. His work includes the development of thermal control systems for spacecraft and the management of research and development activities on thermal control devices.
Keiichi Yanagase is a researcher in mechanical environmental testing and structural engineering
at JAXA. He received his BS and MS degrees in materials engineering from Tohoku University
in 2005 and 2007, respectively, and his PhD in aerospace engineering from Tottori University in
2021.
Kazuki Watanabe is the president and CEO of Well Research Inc. and an inventor of the
Vibration Isolation System (VIS). He earned his BS degree in aerospace engineering from
Tokai University in 1982, his MS degree in mechanical engineering from Waseda University
in 2002, and his PhD from Waseda University in 2005.
Seiji Yoshida is a principal engineer at Sumitomo Heavy Industries, Ltd. He earned both his BS
and MS degrees in mechanical engineering from Nagaoka University of Technology in 1996 and
1998, respectively. His primary focus is on pioneering the development of advanced cryogenic
equipment for space applications.
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