Trajectories
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docs/KNOWLEDGEBASE.md
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# Knowledgebase Guide
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## Overview
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The **Knowledgebase** is a reference library of propellants, materials, and engineering equations. It provides the properties needed for calculations in the Solver, Engine Designer, and Rocket Designer.
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## Structure
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The knowledgebase is organized into four categories:
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1. **Fuels & Oxidizers** — Propellant properties
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2. **Ablative Materials** — Thermal protection
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3. **Structural Materials** — Engine & tank construction
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4. **Equations** — Mathematical references
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---
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## Fuels & Oxidizers
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### What are They?
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**Fuel**: Hydrocarbon or hydrogen source
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- Examples: RP-1 (kerosene), methane, hydrogen
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**Oxidizer**: Oxygen source
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- Examples: Liquid oxygen (LOX), nitrogen tetroxide (N2O4)
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### Key Properties
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Each propellant entry contains:
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**Chemical Properties:**
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- Molecular formula (e.g., O₂ for oxygen)
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- Molecular weight (g/mol)
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- Density (kg/m³)
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**Combustion Properties:**
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- **Adiabatic flame temperature (Tc)** — peak temperature of combustion (K)
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- Higher Tc = higher Isp
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- Example: LOX/RP-1 ≈ 3750 K, LOX/LH2 ≈ 3900 K
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- **Expansion ratio (gamma, γ)** — specific heat ratio
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- γ = Cp / Cv
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- Affects nozzle performance
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- LOX/RP-1: γ ≈ 1.25, LOX/LH2: γ ≈ 1.26
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- **Characteristic velocity (c*)** — ideal nozzle exit velocity
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- Fundamental property of propellant
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- Used to estimate Isp
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**Performance Metrics:**
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- **Specific impulse (vacuum, Isp_vac)** — seconds
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- Higher = better
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- Example: LOX/RP-1 ≈ 310–320 s, LOX/LH2 ≈ 450 s
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- **Specific impulse (sea level, Isp_sl)** — reduced due to ambient pressure
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- Lower than vacuum value
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- Example: LOX/RP-1 ≈ 260 s at sea level
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**Mixture Ratio:**
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- **O/F ratio** — oxidizer mass per fuel mass
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- Example: LOX/RP-1 optimal ≈ 2.5–2.8
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- Higher ratio = more oxygen = higher T but lower mass fraction
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- Lower ratio = more fuel = lower T but better mass fraction
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### Common Propellant Pairs
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| Fuel | Oxidizer | Isp (s) | Tc (K) | γ | Notes |
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|------|----------|---------|-------|-------|-------|
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| RP-1 | LOX | 310 | 3750 | 1.25 | Space-proven, storable |
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| Methane | LOX | 330 | 3900 | 1.24 | Better impulse than RP-1 |
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| Hydrogen | LOX | 450 | 3900 | 1.26 | Best impulse, cryogenic, low density |
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| MMH | N2O4 | 290 | 3400 | 1.20 | Storable, hypergolic (ignites on contact) |
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| Hydrazine | N2O4 | 310 | 3700 | 1.24 | Toxic but storable, high density |
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### Storage Types
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**Cryogenic:**
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- Requires refrigeration
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- LOX, LH2, LN2, LCH4
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- Higher energy density
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- Complex ground support needed
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**Storable (Room Temperature):**
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- RP-1, Methane (marginally), MMH, Hydrazine
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- No cryogenic handling needed
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- Lower energy density than cryo
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- Easier integration
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**Hypergolic:**
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- Ignites spontaneously on contact (fuel + oxidizer)
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- No ignition system needed
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- Toxic, corrosive, more expensive
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- Examples: MMH/N2O4, Hydrazine/N2O4
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### Selection Criteria
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**High Performance** → LOX/LH2 (Isp 450 s, but cryogenic complexity)
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**Best Balance** → LOX/RP-1 (Isp 310 s, storable, proven)
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**Storable Only** → MMH/N2O4 (Isp 290 s, no cryo needed)
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**Simple & Solid** → RP-1/LOX or Methane/LOX
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---
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## Ablative Materials
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### Purpose
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Ablative materials line the rocket engine chamber to withstand:
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- **High temperature** (combustion products: 3500+ K)
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- **High pressure** (200+ bar)
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- **Erosion** from hot gas flow
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### Key Properties
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Each ablative material entry contains:
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**Mechanical Properties:**
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- Density (kg/m³)
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- Tensile strength (MPa)
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- Compressive strength (MPa)
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**Thermal Properties:**
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- Specific heat (J/kg·K)
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- Thermal conductivity (W/m·K)
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- Melting point (K)
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**Erosion Properties:**
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- **Base erosion rate** (inch/s @ 300 psi reference pressure)
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- Ablative material mass removed per unit time
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- Higher rate = thicker liner needed
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- Example: PAXS ≈ 0.025 in/s
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- **Pressure exponent (n)** — power-law sensitivity to pressure
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- `rate(P) = base_rate × (P / P_ref)^n`
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- Lower n = less pressure-sensitive (better for high P)
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- Example: PAXS n = 0.38, KFSI n = 0.35
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**Application Notes:**
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- Where used: chamber, nozzle, injector
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- Temperature limits
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- Compatibility with propellants
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### Material Classes
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**Composites (Rigid):**
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- PAXS (polyester + glass) — n = 0.38
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- Common, cost-effective
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- Moderate erosion rate
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- Max temp ~2000 K
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- KFSI (silica + phenolic) — n = 0.35
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- Better than PAXS
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- Lower erosion rate
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- Higher temp capability
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- More expensive
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- Carbon-Phenolic — n = 0.32
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- Excellent thermal performance
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- Very low erosion rate
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- Very expensive
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- Used on high-end engines
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**Elastomers (Flexible):**
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- ZIRCONIA (zirconia + silicone) — n = 0.48
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- Flexible (less brittle)
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- Higher erosion rate
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- Absorbs vibration
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- Lower cost than phenolics
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- Butyl Rubber — n = 0.50
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- Very flexible
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- High erosion rate
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- Used for low-pressure applications
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### Pressure Exponent Explanation
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The power law `rate ∝ P^n` comes from **regression rate burning**:
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- Composites: n ≈ 0.3–0.4
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- Erosion driven by thermal decomposition
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- Less pressure-dependent
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- Preferred for high-pressure engines
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- Elastomers: n ≈ 0.4–0.5
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- Erosion driven by shear stress
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- More pressure-dependent
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- OK for moderate pressures
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**High Pressure (>1000 psi):** Use composite with low n
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**Moderate Pressure (500 psi):** Elastomer acceptable
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**Low Pressure (<300 psi):** Elastomer preferred (lower cost)
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### Selection Guidance
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| Application | Material | Reason |
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|-------------|----------|--------|
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| High-perf liquid rocket | KFSI or Carbon-Phenolic | Low erosion rate, high pressure capable |
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| Medium-perf liquid | PAXS | Good balance, cost-effective |
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| Hybrid rocket | KFSI (fuel grain liner) | Low ablation to preserve geometry |
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| Solid rocket | Composite | Erosion from particles & hot gas |
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| Low-cost experiment | Butyl or PAXS | Adequate performance, lowest cost |
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---
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## Structural Materials
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### Purpose
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Structural materials form the engine chamber and rocket tanks. They must:
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- Withstand internal pressure (hoop stress)
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- Resist corrosion from propellants
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- Perform at operating temperature
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- Balance weight vs. strength
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### Key Properties
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Each structural material entry contains:
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**Mechanical Properties:**
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- **Yield strength** (MPa) — stress at which permanent deformation begins
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- Higher = thinner walls possible = lighter
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- Example: Aluminum 6061-T6: 275 MPa, Inconel 718: 1240 MPa
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- **Young's modulus** (GPa) — stiffness
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- Higher = less deflection (stronger)
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- **Density** (kg/m³) — weight per volume
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- Lower = lighter vehicle
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- Example: Al: 2700, Ti: 4430, SS: 8000
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**Thermal Properties:**
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- Melting point (K) — temperature limit
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- Coefficient of thermal expansion (CTE)
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- Thermal conductivity
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**Other:**
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- Cost (relative)
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- Machinability
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- Availability
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### Material Comparison
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| Material | Density | Yield | T_melt | Cost | Best For |
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|----------|---------|-------|--------|------|----------|
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| Al 6061-T6 | 2700 | 275 | 933 | 1× | Prototype, pressure-fed |
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| SS 304 | 8000 | 215 | 1726 | 3× | Corrosion resistance |
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| Inconel 718 | 8190 | 1240 | 1600 | 10× | High-performance engines |
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| Ti-6-4 | 4430 | 880 | 1941 | 15× | Lightweight, space |
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| CFRP | 1600 | 600+ | 700 | 20× | Lowest weight |
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### Hoop Stress Calculation
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For cylindrical pressure vessels:
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```
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σ_hoop = (P × r) / t
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```
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Setting equal to yield with safety factor:
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```
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t = (P × r) / (σ_yield / SF)
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```
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Higher yield strength → thinner walls → lighter mass
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### Material Selection
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**Pressure-Fed Engine (200+ bar, chamber):**
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- **Inconel 718** — withstands high pressure & heat
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- **Titanium** — lighter alternative
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- **Aluminum** — cheap but needs cooling design
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**Rocket Tanks (20–50 bar):**
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- **Aluminum 6061** — standard, proven, affordable
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- **Titanium** — if weight critical
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- **CFRP** — lowest weight, highest cost
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**Low-Pressure Systems:**
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- **Aluminum** — sufficient, cheapest
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- **Stainless Steel** — if corrosion concern
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---
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## Equations Reference
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### Fundamental Relations
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**Thrust (momentum equation):**
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```
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F = ṁ · Ve + (Pe - Pa) · Ae
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```
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- ṁ = mass flow rate
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- Ve = exit velocity
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- Pe = exit pressure
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- Pa = ambient pressure
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- Ae = exit area
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**Specific Impulse:**
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```
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Isp = Ve / g0 (vacuum)
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Isp = (Ve - (Pe - Pa) / ṁ · Ae) / g0 (sea level)
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```
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- g0 = 9.81 m/s² (gravitational acceleration)
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**Rocket Equation (Tsiolkovsky):**
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```
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ΔV = Isp · g0 · ln(m_initial / m_final)
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```
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- ΔV = velocity change (delta-v)
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- m_initial = wet mass
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- m_final = dry mass
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### Isentropic Flow (Nozzle Design)
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**Temperature at exit:**
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```
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Te / T0 = (Pe / P0)^((γ-1)/γ)
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```
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**Density at exit:**
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```
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ρe / ρ0 = (Pe / P0)^(1/γ)
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```
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**Mach number from area ratio:**
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```
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A / A* = (1/M) · [(2/(γ+1)) · (1 + (γ-1)/2 · M²)]^((γ+1)/(2(γ-1)))
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```
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- A* = throat area
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- A = arbitrary section area
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- M = Mach number
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- (requires numerical inversion via bisection)
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**Characteristic velocity:**
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```
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c* = √[2 · (γ+1)/(γ-1) · R · T0 · (2/(γ+1))^((γ+1)/(γ-1))]
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```
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### Thrust Coefficient
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```
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CF = √[2γ²/(γ-1) · (2/(γ+1))^((γ+1)/(γ-1)) · (1 - (Pe/P0)^((γ-1)/γ))]
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+ (Pe - Pa)/(P0) · (Ae/At)
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```
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### Chamber Thermodynamics
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**Energy balance (no losses):**
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```
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Cp · (T0 - Te) = Ve² / 2
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```
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- Cp = specific heat at constant pressure
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**Chemical equilibrium:**
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Determine T0, γ from propellant properties and stoichiometry
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(computed via NASA CEA code or thermodynamic tables)
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### Drag & Atmosphere
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**Drag force:**
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```
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Fd = 0.5 · ρ · v² · Cd · A
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```
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- ρ = air density (depends on altitude)
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- v = velocity
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- Cd = drag coefficient (~0.25 for rockets)
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- A = reference area
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**US Standard Atmosphere (piecewise):**
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```
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T(h) = T0 - L·h (troposphere, 0–11 km)
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P(h) = P0 · (T(h)/T0)^(-g/(R·L))
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ρ(h) = ρ0 · (T(h)/T0)^(-(g/(R·L) + 1))
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```
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- L = lapse rate ≈ 6.5 K/km
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- R = gas constant
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### Hoop Stress (Pressure Vessels)
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**Thin-walled cylinder:**
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```
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σ_hoop = P·r / t
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```
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With safety factor:
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```
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t = P·r / (σ_yield / SF)
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```
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Mass:
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```
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m = 2π · r · t · L · ρ (cylinder)
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m = 4π · r² · t · ρ (hemispherical dome)
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```
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### Mass Fraction
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**Payload fraction:**
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```
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fp = m_payload / m_wet
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```
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**Structure fraction:**
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```
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fs = m_structure / m_wet
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```
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**Propellant fraction:**
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```
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fp = m_propellant / m_wet
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```
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Sum: `fp + fs + fpayload = 1`
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Typical rockets: `fs = 0.10–0.20` (10–20%)
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---
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## Using Knowledgebase in Design
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### Importing Propellant Properties
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1. **Solver**: Drag `chamberTemperature`, `expansionGamma` onto workspace
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2. Check knowledgebase for propellant pair (LOX + RP-1)
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3. Enter T0 from knowledgebase → solver computes Isp
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4. Verify with reference Isp in database
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### Cross-Checking Ablation
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1. **Engine Design**: Select PAXS ablative
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2. **Check Pressure Exponent**: 0.38 (from knowledgebase)
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3. Verify erosion rate correction is applied
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4. Compare remaining thickness to safety margin (>0.5 inch)
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### Material Trade-Studies
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1. **Rocket Design**: Compare materials
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- Aluminum: 1000 kg tanks
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- Titanium: 600 kg tanks (lighter, more expensive)
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- CFRP: 350 kg tanks (lightest, most expensive)
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2. Use mass to compute TWR, delta-v
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3. Pick best balance for mission
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---
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## Future Enhancements
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- **User contributions**: Allow adding new materials/propellants
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- **Temperature corrections**: Adjust properties with temperature
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- **Material compatibility**: Show fuel/material interactions
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- **Cost database**: Include material & propellant costs
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- **References**: Link to papers & technical data sources
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- **Lookup plots**: Interactive charts (Isp vs. O/F, etc.)
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---
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## References
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- NASA SP-273: Liquid Rocket Engine Combustion Instability.
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- Huzel, D. K., & Huang, D. H. (1992). Modern engineering for design of liquid-propellant rocket engines. AIAA.
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- Rocket Propulsion Elements (8th ed.). Sutton & Biblarz.
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- US Standard Atmosphere 1976. NASA TM-X-74335.
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- CEA (Chemical Equilibrium with Applications). NASA Glenn Research Center.
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---
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**Last Updated**: 2025-02 | **Status**: Current
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