Design and Experimental Validation of the Separation and Deployment System for the CHERI Lunar Micro-Rovers

Ahmet Mustafa Kangal; Oğuzhan Şahan; Teoman Efe Aycan; Görkem Egemen Güloğlu; Halil Ersin Söken; Ali Emre Turgut; Uluç Saranlı; Mustafa Mert Ankaralı

cs.RONov 2, 2026

Design and Experimental Validation of the Separation and Deployment System for the CHERI Lunar Micro-Rovers

Authors: Ahmet Mustafa Kangal, Oğuzhan Şahan, Teoman Efe Aycan, Görkem Egemen Güloğlu, Halil Ersin Söken, Ali Emre Turgut, Uluç Saranlı, Mustafa Mert Ankaralı

Organizations: Robotics Graduate Program, Graduate School of Natural and Applied Sciences, Middle East Technical University (METU), Ankara, Türkiye · Department of Aerospace Engineering, Middle East Technical University (METU), Ankara, Türkiye

Abstract

Reliable separation mechanisms are critical for planetary micro-rovers, which must withstand launch vibrations and then eject the vehicle clear of the lander. This paper presents the design, optimization, and testing of the separation system of the CHERI (Challenging Environment Exploration Rovers for Intelligence) lunar micro-rovers. Sized for a 5 kg vehicle, the assembly comprises an aluminum ground plate, a central non-explosive hold-down and release mechanism, and an asymmetric four-spring ejection array. Initial separation trials revealed a velocity deficit caused by energy absorption in the flexible wheel-legs and chassis during the ejection stroke. This loss was represented as equivalent chassis damping in a multi-body dynamic model, and the springs were re-optimized accordingly. Early hardware iterations failed vibration testing, first through spring jamming in the guide housings and then through a 34 Hz fundamental mode of the flat ground plate. A density-based topology optimization raised the fundamental frequency to 84 Hz, above the 70 Hz requirement of the lander, within a 1.6 kg mass budget. The final configuration passed acceptance-level sinusoidal vibration, random vibration, and shock testing without structural degradation. Separation tests with a weight-matched one-sixth mass model, tracked by optical motion capture, confirmed clearance of the lander and reproduced the simulated free flight to within the run-to-run variation of the measured ejection velocity. These tests further showed that landing outcome is dominated by touchdown and settling dynamics, indicating that the residual pitch rate is the main controllable design parameter and confirming self-righting recovery as an integral element of the deployment architecture.

Figures & tables

Figure 1: Concept illustration of the CHERI micro-rovers (AI-generated image, shown for illustration only).

Panel of Figure 2 — Figure 2: The CHERI micro-rover platform: (a) CAD render of the rover; (b) orthographic technical drawing indicating key structural dimensions and the hold-down interface; (c) the manufactured Structural Model (SM) used for mechanical verification and environmental testing.

Panel of Figure 3 — Figure 3: Free-fall deployment kinematics showing the stowed vertical mounting height ( $H=1.7\,\text{m}$ ) and the minimum dynamic horizontal clearance ( $D_{\text{clearance}}\geq 0.7\,\text{m}$ ) required to clear the host lander’s landing pads.

Figure 4: Labeled exploded assembly view of the CHERI micro-rover separation interface, detailing the key physical components and sequence of structural integration: (0) lander surface, (1) isolation pads, (2) ground plate, (3) mounting screws, (4) separation springs, (5) separation nut, and (6) CHERI Rover.

Panel of Figure 5 — Figure 5: First technological era using micro shape memory alloy (SMA) release nuts (Iterations 1–3): (a–c) isometric CAD models; (d–f) manufactured hardware prototypes.

Panel of Figure 6 — Figure 6: Second technological era using a fracture-type separation nut 22 (Iterations 4–6): (a–c) isometric CAD models; (d–f) corresponding physical hardware.

Iter.	Actuator Type	Spring Config.	Mass (kg)	1st Nat. Freq.	Critical Failure Mode	Status
1	2 $\times$ Release Nut	None (Gravity)	0.85	NA	Actuator synchronization; no ejection force	NA
2	1 $\times$ Release Nut	2 $\times$ Lateral	0.92	NA	Insufficient clearance (clips lander footpad)	NA
3	1 $\times$ Release Nut	2 $\times$ Vertical	1.05	NA	Inadequacy under cyclic test conditions	NA
4	1 $\times$ Separation Nut	4 $\times$ Asymmetric	1.28	51 Hz	Spring stiction; spring jam in Z-axis random vib.	Failed
5	1 $\times$ Separation Nut	4 $\times$ Asym. (Opt. Guides)	1.48	34 Hz	First mode below lander requirement (70 Hz )	Failed
6	1 $\times$ Separation Nut	4 $\times$ Asym. (Opt. Guides)	1.54	84 Hz	None (SIMP topology-optimized plate)	Passed

Table 1: System-level engineering comparison of the six evolutionary separation system generations.

Figure 7: Chronological evolution of the separation spring mechanisms, detailing CAD cross-sections (top row) and corresponding physical hardware (bottom row). Left (Phase A): the initial enclosed COTS-based baseline. Middle (Phase B): the first custom iteration, whose unoptimized guide allowed lateral buckling and sliding stiction during release. Right (Phase C): the final configuration with the re-engineered guide bushing and rod. Key components are annotated: (1) spring hat, (2) helical spring, (3) guide bushing, and (4) central guide rod. The re-engineered guide bushing (3) and rod (4) ensure a stable, linear stroke without buckling or stiction.

Panel of Figure 8 — Figure 8: Two-stage parametric spring optimization: (a) increasing the spring rate by $+53\%$ while keeping the energy delivered per spring over the separation stroke constant at $680.3\,\text{mJ}$ ; (b) increasing the free length from $82.0\,\text{mm}$ (Proposed) to $85.0\,\text{mm}$ (Optimum), which raises the spring force at the release point ( $75.0\,\text{mm}$ ) from $\approx 9.8\,\text{N}$ to $\approx 12.1\,\text{N}$ with a similar stowed force.

Config.	Free length	Rate	Coils	Rate incr.
	$L_{f}$ (mm)	$k$ (N/mm)	$n_{t}$	(%)
Original	92.4	0.910	16.0	–
Proposed (82)	82.0	1.396	11.1	+53.4%
Optimized (85)	85.0	1.209	12.5	+32.9%

Table 2: Geometric and physical comparison of the evolutionary compression spring iterations.

Panel of Figure 9 — Figure 9: Finite element modal analysis comparison showing the first fundamental mode shapes of (a) the flat fifth-generation ground plate exhibiting severe central flexing at $34\,\text{Hz}$ , and (b) the topology-optimized sixth-generation ground plate with stiffening ribs raising the natural frequency to $84\,\text{Hz}$ .

Mode	Deformation Shape	Pre-Opt.	Post-Opt.	Status
No.		Iter. 5 (Hz)	Iter. 6 (Hz)	( $\geq 70\,\text{Hz}$ )
1	Longitudinal Torsion	33.6	84.0	Passed
2	Out-of-plane Bending	55.4	116.9	Passed
3	Lateral Torsion	59.8	107.2	Passed

Table 3: First three natural frequencies of the stowed separation assembly before and after topology optimization (FEA). Modes are matched by deformation shape.

Panel of Figure 10 — Figure 10: Structural test and calibration setups at the TÜBİTAK UZAY satellite testing facility: (a) the stowed separation assembly mounted on the electrodynamic shaker table for vibration testing in three orthogonal axes, (b) the pneumatic shock test setup, in which an aluminum projectile strikes a resonant plate, and (c) the non-functional dummy model used as a mass simulator.

Quantity	Scaling law	Ratio (Earth/Moon)
Weight	$f\,(g_{0}/g_{\text{Moon}})$	$1.01$
Downrange at touchdown	$f^{-1/2}(g_{\text{Moon}}/g_{0})^{1/2}$	$1.00$
Attitude at touchdown	$f^{-1/2}(g_{\text{Moon}}/g_{0})^{1/2}$	$1.00$
Translational kinetic energy	$f\cdot f^{-1}$	$1.00$
Rotational kinetic energy	$f\cdot f^{-1}$	$1.00$
Separation velocity	$f^{-1/2}$	$2.45$

Table 4: Ratio of the terrestrial

1/6

mass test to lunar flight conditions under weight-matched Froude scaling, with

f=m_{\text{sim}}/m_{\text{flight}}=1/6

and

g_{0}/g_{\text{Moon}}=6.06

. Quantities with unit ratio transfer exactly; the remainder do not.

Figure 11: Physical experimental launch test setup, showing the test arena with wall-mounted optical motion-capture cameras and the landing sandbox (left), and the $1/6$ mass model stowed and preloaded on the ground plate (right).

Figure 12: Terrestrial test rig at three representative slope conditions used in the campaign of Section 6.3 : (left to right) $0^{\circ}$ , $+10^{\circ}$ and $+20^{\circ}$ . Steeper and clockwise ( $-20^{\circ}$ and beyond) conditions are reported in Table 5 but not shown here, as they were tested with fewer repetitions after early overturns (Section 6.3 ).

Slope (deg)	Lunar sim.	Earth sim.	Earth exp.
$-45$	$7/10$	$0/10$	$0/5$
$-30$	$10/10$	$2/10$	$0/5$
$-20$	$10/10$	$\mathbf{0/10}$	$0/5$
$-10$	$10/10$	$8/10$	$3/10$
$0$	$10/10$	$8/10$	$4/10$
$+10$	$10/10$	$9/10$	$5/10$

Table 5: Separation and landing outcomes (successes/runs) over the slope range. Lunar simulation:

5.0\,\text{kg}

g=1.62\,\text{m/s}^{2}

. Terrestrial simulation and experiment: both

0.833\,\text{kg}

g=9.81\,\text{m/s}^{2}

. Positive slopes are counter-clockwise, negative clockwise. Cells with

5

runs were truncated after repeated failure to conserve separation-nut actuations.

Figure 13: Motion-capture trajectory of the $1/6$ mass model during separation at $+20^{\circ}$ slope (right), compared with the corresponding terrestrial simulation (left). Gray markers denote interpolated positions where tracking data were unavailable.

Figure 14: Motion-capture trajectory of the $1/6$ mass model during separation at $-20^{\circ}$ slope (right), compared with the corresponding terrestrial simulation (left).

Figure 15: Simulated landing outcomes on a $-20^{\circ}$ slope. Left: lunar conditions, full flight mass, successful wheels-first landing. Right: terrestrial conditions, $1/6$ mass model, overturned. The free-flight trajectories are geometrically identical; the divergence occurs entirely during touchdown and settling.

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