Propulsion Assessment Skill

Read CONVENTIONS.md at the repo root before proceeding.

This skill sizes the propulsion subsystem hardware — engines, tanks, and feed systems. It consumes the Delta-V budget from mission-analysis-specialist and produces propellant mass and hardware specifications for the mass/power budgets.

Before You Begin

Ask the user (if not already known):

1. What is the mission type and target body?
2. Chemical or electric propulsion? (or both — many missions use chemical for large burns and EP for station keeping)
3. Are there constraints on propellant type? (e.g., "green" non-toxic, heritage hydrazine, bipropellant, cold gas for CubeSats)
4. What standards framework applies?
5. What design phase?

Applicable Phases

• Primary: Phase A (trade studies, first-order sizing), Phase B (preliminary design)
• Supporting: Phase C (performance verification), Phase D (propellant loading planning)

Ownership Boundary

| Responsibility | Owner | |:---|:---| | Delta-V budget (maneuver list, margins) | mission-analysis-specialist | | Propellant mass, engine selection, tank sizing, feed system | This skill | | Mass budget integration | systems-engineering-assessment |

Analysis Workflows

1. Propellant Mass Sizing (Chemical)

• Input: Total Delta-V (from mission-analysis-specialist), spacecraft dry mass, engine Isp.
• Tsiolkovsky Equation: $m_p = m_{dry} \cdot (e^{\Delta V / (I_{sp} \cdot g_0)} - 1)$
• Margin: Add ullage (typically 3-5%), residuals (1-2%), and loading uncertainty (1%).
• Reference Isp values (ask user to confirm or provide actual):
  • Cold gas (N₂): 65-70s
  • Monopropellant (Hydrazine): 220-230s
  • Green monoprop (AF-M315E/LMP-103S): 235-255s
  • Bipropellant (MMH/NTO): 310-320s
  • Bipropellant (LOX/LH2): 440-460s

2. Electric Propulsion Sizing

• Key difference: EP is power-limited, not propellant-limited. Thrust is low, burn durations are weeks to months.
• Input: Delta-V, available power from power-assessment, mission timeline constraints.
• Thrust: $F = 2 \eta P / (g_0 \cdot I_{sp})$ where $\eta$ is thruster efficiency.
• Burn duration: $t_b = m_p \cdot I_{sp} \cdot g_0 / F$
• Reference Isp values:
  • Electrospray: 800-1500s
  • Hall-effect: 1200-1800s
  • Gridded ion (e.g., NSTAR): 2500-3500s
• Power requirement: Typically 1-30 kW depending on thruster type. Flag if this exceeds the power budget.

3. Tank Sizing

• Volume: $V = m_p / \rho_{prop}$ with ullage margin (typically 5-10%).
• Material: Titanium (heritage), COPV (mass savings), aluminum (budget).
• Pressure: MEOP for blowdown vs. regulated systems.

4. Thrust-to-Weight (Chemical Only)

• Powered descent/ascent: Require $T/W > 1.0$ relative to local gravity.
• Orbit maneuvers: T/W is less critical but affects gravity losses (flag if $T/W < 0.1$).

Output Format

1. Propulsion Report (propulsion_report.md):
   • Engine selection rationale (type, Isp, thrust level)
   • Propellant mass breakdown (usable, ullage, residual, total)
   • Tank specifications (volume, pressure, material)
   • Burn durations per maneuver
   • 🟢 / 🟡 / 🔴 feasibility status
2. Feasibility flags: Alert if propellant mass exceeds reasonable wet-mass fraction (>40% for most missions, >80% for lunar/interplanetary)

Interface

• Reads from: /requirements/, /analysis/mission-analysis-specialist/ (Delta-V budget), /analysis/power-assessment/ (for EP power availability)
• Writes to: /analysis/propulsion-assessment/