rocketry/docs/KNOWLEDGEBASE.md

Knowledgebase Guide

Overview

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.

Structure

The knowledgebase is organized into four categories:

1. Fuels & Oxidizers — Propellant properties
2. Ablative Materials — Thermal protection
3. Structural Materials — Engine & tank construction
4. Equations — Mathematical references

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Fuels & Oxidizers

What are They?

Fuel: Hydrocarbon or hydrogen source

• Examples: RP-1 (kerosene), methane, hydrogen

Oxidizer: Oxygen source

• Examples: Liquid oxygen (LOX), nitrogen tetroxide (N2O4)

Key Properties

Each propellant entry contains:

Chemical Properties:

• Molecular formula (e.g., O₂ for oxygen)
• Molecular weight (g/mol)
• Density (kg/m³)

Combustion Properties:

• Adiabatic flame temperature (Tc) — peak temperature of combustion (K)

  • Higher Tc = higher Isp
  • Example: LOX/RP-1 ≈ 3750 K, LOX/LH2 ≈ 3900 K

• Expansion ratio (gamma, γ) — specific heat ratio

  • γ = Cp / Cv
  • Affects nozzle performance
  • LOX/RP-1: γ ≈ 1.25, LOX/LH2: γ ≈ 1.26

• Characteristic velocity (c)* — ideal nozzle exit velocity

  • Fundamental property of propellant
  • Used to estimate Isp

Performance Metrics:

• Specific impulse (vacuum, Isp_vac) — seconds

  • Higher = better
  • Example: LOX/RP-1 ≈ 310–320 s, LOX/LH2 ≈ 450 s

• Specific impulse (sea level, Isp_sl) — reduced due to ambient pressure

  • Lower than vacuum value
  • Example: LOX/RP-1 ≈ 260 s at sea level

Mixture Ratio:

• O/F ratio — oxidizer mass per fuel mass
  • Example: LOX/RP-1 optimal ≈ 2.5–2.8
  • Higher ratio = more oxygen = higher T but lower mass fraction
  • Lower ratio = more fuel = lower T but better mass fraction

Common Propellant Pairs

Fuel	Oxidizer	Isp (s)	Tc (K)	γ	Notes
RP-1	LOX	310	3750	1.25	Space-proven, storable
Methane	LOX	330	3900	1.24	Better impulse than RP-1
Hydrogen	LOX	450	3900	1.26	Best impulse, cryogenic, low density
MMH	N2O4	290	3400	1.20	Storable, hypergolic (ignites on contact)
Hydrazine	N2O4	310	3700	1.24	Toxic but storable, high density

Storage Types

Cryogenic:

• Requires refrigeration
• LOX, LH2, LN2, LCH4
• Higher energy density
• Complex ground support needed

Storable (Room Temperature):

• RP-1, Methane (marginally), MMH, Hydrazine
• No cryogenic handling needed
• Lower energy density than cryo
• Easier integration

Hypergolic:

• Ignites spontaneously on contact (fuel + oxidizer)
• No ignition system needed
• Toxic, corrosive, more expensive
• Examples: MMH/N2O4, Hydrazine/N2O4

Selection Criteria

High Performance → LOX/LH2 (Isp 450 s, but cryogenic complexity) Best Balance → LOX/RP-1 (Isp 310 s, storable, proven) Storable Only → MMH/N2O4 (Isp 290 s, no cryo needed) Simple & Solid → RP-1/LOX or Methane/LOX

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Ablative Materials

Purpose

Ablative materials line the rocket engine chamber to withstand:

• High temperature (combustion products: 3500+ K)
• High pressure (200+ bar)
• Erosion from hot gas flow

Key Properties

Each ablative material entry contains:

Mechanical Properties:

• Density (kg/m³)
• Tensile strength (MPa)
• Compressive strength (MPa)

Thermal Properties:

• Specific heat (J/kg·K)
• Thermal conductivity (W/m·K)
• Melting point (K)

Erosion Properties:

• Base erosion rate (inch/s @ 300 psi reference pressure)

  • Ablative material mass removed per unit time
  • Higher rate = thicker liner needed
  • Example: PAXS ≈ 0.025 in/s

• Pressure exponent (n) — power-law sensitivity to pressure

  • rate(P) = base_rate × (P / P_ref)^n
  • Lower n = less pressure-sensitive (better for high P)
  • Example: PAXS n = 0.38, KFSI n = 0.35

Application Notes:

• Where used: chamber, nozzle, injector
• Temperature limits
• Compatibility with propellants

Material Classes

Composites (Rigid):

• PAXS (polyester + glass) — n = 0.38

  • Common, cost-effective
  • Moderate erosion rate
  • Max temp ~2000 K

• KFSI (silica + phenolic) — n = 0.35

  • Better than PAXS
  • Lower erosion rate
  • Higher temp capability
  • More expensive

• Carbon-Phenolic — n = 0.32

  • Excellent thermal performance
  • Very low erosion rate
  • Very expensive
  • Used on high-end engines

Elastomers (Flexible):

• ZIRCONIA (zirconia + silicone) — n = 0.48

  • Flexible (less brittle)
  • Higher erosion rate
  • Absorbs vibration
  • Lower cost than phenolics

• Butyl Rubber — n = 0.50

  • Very flexible
  • High erosion rate
  • Used for low-pressure applications

Pressure Exponent Explanation

The power law rate ∝ P^n comes from regression rate burning:

• Composites: n ≈ 0.3–0.4

  • Erosion driven by thermal decomposition
  • Less pressure-dependent
  • Preferred for high-pressure engines

• Elastomers: n ≈ 0.4–0.5

  • Erosion driven by shear stress
  • More pressure-dependent
  • OK for moderate pressures

High Pressure (>1000 psi): Use composite with low n Moderate Pressure (500 psi): Elastomer acceptable Low Pressure (<300 psi): Elastomer preferred (lower cost)

Selection Guidance

Application	Material	Reason
High-perf liquid rocket	KFSI or Carbon-Phenolic	Low erosion rate, high pressure capable
Medium-perf liquid	PAXS	Good balance, cost-effective
Hybrid rocket	KFSI (fuel grain liner)	Low ablation to preserve geometry
Solid rocket	Composite	Erosion from particles & hot gas
Low-cost experiment	Butyl or PAXS	Adequate performance, lowest cost

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Structural Materials

Purpose

Structural materials form the engine chamber and rocket tanks. They must:

• Withstand internal pressure (hoop stress)
• Resist corrosion from propellants
• Perform at operating temperature
• Balance weight vs. strength

Key Properties

Each structural material entry contains:

Mechanical Properties:

• Yield strength (MPa) — stress at which permanent deformation begins

  • Higher = thinner walls possible = lighter
  • Example: Aluminum 6061-T6: 275 MPa, Inconel 718: 1240 MPa

• Young's modulus (GPa) — stiffness

  • Higher = less deflection (stronger)

• Density (kg/m³) — weight per volume

  • Lower = lighter vehicle
  • Example: Al: 2700, Ti: 4430, SS: 8000

Thermal Properties:

• Melting point (K) — temperature limit
• Coefficient of thermal expansion (CTE)
• Thermal conductivity

Other:

• Cost (relative)
• Machinability
• Availability

Material Comparison

Material	Density	Yield	T_melt	Cost	Best For
Al 6061-T6	2700	275	933	1×	Prototype, pressure-fed
SS 304	8000	215	1726	3×	Corrosion resistance
Inconel 718	8190	1240	1600	10×	High-performance engines
Ti-6-4	4430	880	1941	15×	Lightweight, space
CFRP	1600	600+	700	20×	Lowest weight

Hoop Stress Calculation

For cylindrical pressure vessels:

σ_hoop = (P × r) / t

Setting equal to yield with safety factor:

t = (P × r) / (σ_yield / SF)

Higher yield strength → thinner walls → lighter mass

Material Selection

Pressure-Fed Engine (200+ bar, chamber):

• Inconel 718 — withstands high pressure & heat
• Titanium — lighter alternative
• Aluminum — cheap but needs cooling design

Rocket Tanks (20–50 bar):

• Aluminum 6061 — standard, proven, affordable
• Titanium — if weight critical
• CFRP — lowest weight, highest cost

Low-Pressure Systems:

• Aluminum — sufficient, cheapest
• Stainless Steel — if corrosion concern

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Equations Reference

Fundamental Relations

Thrust (momentum equation):

F = ṁ · Ve + (Pe - Pa) · Ae

• ṁ = mass flow rate
• Ve = exit velocity
• Pe = exit pressure
• Pa = ambient pressure
• Ae = exit area

Specific Impulse:

Isp = Ve / g0  (vacuum)
Isp = (Ve - (Pe - Pa) / ṁ · Ae) / g0  (sea level)

• g0 = 9.81 m/s² (gravitational acceleration)

Rocket Equation (Tsiolkovsky):

ΔV = Isp · g0 · ln(m_initial / m_final)

• ΔV = velocity change (delta-v)
• m_initial = wet mass
• m_final = dry mass

Isentropic Flow (Nozzle Design)

Temperature at exit:

Te / T0 = (Pe / P0)^((γ-1)/γ)

Density at exit:

ρe / ρ0 = (Pe / P0)^(1/γ)

Mach number from area ratio:

A / A* = (1/M) · [(2/(γ+1)) · (1 + (γ-1)/2 · M²)]^((γ+1)/(2(γ-1)))

• A* = throat area
• A = arbitrary section area
• M = Mach number
• (requires numerical inversion via bisection)

Characteristic velocity:

c* = √[2 · (γ+1)/(γ-1) · R · T0 · (2/(γ+1))^((γ+1)/(γ-1))]

Thrust Coefficient

CF = √[2γ²/(γ-1) · (2/(γ+1))^((γ+1)/(γ-1)) · (1 - (Pe/P0)^((γ-1)/γ))]
    + (Pe - Pa)/(P0) · (Ae/At)

Chamber Thermodynamics

Energy balance (no losses):

Cp · (T0 - Te) = Ve² / 2

• Cp = specific heat at constant pressure

Chemical equilibrium: Determine T0, γ from propellant properties and stoichiometry (computed via NASA CEA code or thermodynamic tables)

Drag & Atmosphere

Drag force:

Fd = 0.5 · ρ · v² · Cd · A

• ρ = air density (depends on altitude)
• v = velocity
• Cd = drag coefficient (~0.25 for rockets)
• A = reference area

US Standard Atmosphere (piecewise):

T(h) = T0 - L·h  (troposphere, 0–11 km)
P(h) = P0 · (T(h)/T0)^(-g/(R·L))
ρ(h) = ρ0 · (T(h)/T0)^(-(g/(R·L) + 1))

• L = lapse rate ≈ 6.5 K/km
• R = gas constant

Hoop Stress (Pressure Vessels)

Thin-walled cylinder:

σ_hoop = P·r / t

With safety factor:

t = P·r / (σ_yield / SF)

Mass:

m = 2π · r · t · L · ρ  (cylinder)
m = 4π · r² · t · ρ  (hemispherical dome)

Mass Fraction

Payload fraction:

fp = m_payload / m_wet

Structure fraction:

fs = m_structure / m_wet

Propellant fraction:

fp = m_propellant / m_wet

Sum: fp + fs + fpayload = 1

Typical rockets: fs = 0.10–0.20 (10–20%)

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Using Knowledgebase in Design

Importing Propellant Properties

1. Solver: Drag chamberTemperature, expansionGamma onto workspace
2. Check knowledgebase for propellant pair (LOX + RP-1)
3. Enter T0 from knowledgebase → solver computes Isp
4. Verify with reference Isp in database

Cross-Checking Ablation

1. Engine Design: Select PAXS ablative
2. Check Pressure Exponent: 0.38 (from knowledgebase)
3. Verify erosion rate correction is applied
4. Compare remaining thickness to safety margin (>0.5 inch)

Material Trade-Studies

1. Rocket Design: Compare materials
   • Aluminum: 1000 kg tanks
   • Titanium: 600 kg tanks (lighter, more expensive)
   • CFRP: 350 kg tanks (lightest, most expensive)
2. Use mass to compute TWR, delta-v
3. Pick best balance for mission

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Future Enhancements

• User contributions: Allow adding new materials/propellants
• Temperature corrections: Adjust properties with temperature
• Material compatibility: Show fuel/material interactions
• Cost database: Include material & propellant costs
• References: Link to papers & technical data sources
• Lookup plots: Interactive charts (Isp vs. O/F, etc.)

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References

• NASA SP-273: Liquid Rocket Engine Combustion Instability.
• Huzel, D. K., & Huang, D. H. (1992). Modern engineering for design of liquid-propellant rocket engines. AIAA.
• Rocket Propulsion Elements (8th ed.). Sutton & Biblarz.
• US Standard Atmosphere 1976. NASA TM-X-74335.
• CEA (Chemical Equilibrium with Applications). NASA Glenn Research Center.

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Last Updated: 2025-02 | Status: Current