Subscale Liquid Rocket Engine Stand (LOX/RP-1)
Regeneratively cooled copper-alloy thrust chamber with coaxial swirl injector head.
Aerospace & Aeronautical Engineering / Liquid Rocket Thrust Chamber
FREE TO EXPLORELABORATORY 33 · AEROSPACE & AERONAUTICAL ENGINEERING
Liquid Rocket Engine Thrust Chamber and Characteristic Velocity
Measure propellant mass flow rate m_dot, chamber pressure P_c, and thrust F to determine characteristic exhaust velocity c* and thrust coefficient C_F. Explore interactive wind tunnel simulation, flight dynamics, and aerospace engineering telemetry.
AERODYNAMICS · PROPULSION · FLIGHT DYNAMICS
Your virtual aerospace engineering laboratory.
CHAMBER PRESSURE P_C—
FLIGHT & TEST CELL TELEMETRY
Ready to sample with m_dot = 3.0 kg/s.
Adjust m_dot, then execute aerospace trial.
Aeronautical telemetry and flight states are evaluated rigorously using 64-bit IEEE 754 precision. Residuals reflect ordinary least-squares regression across successive observations. Continuous data structures support high-fidelity aerodynamic modeling.
Aeronautical assumptions and boundary condition limits
Regenerative coolant heat pick-up increases propellant injection temperature, slightly modifying theoretical c*. Linear regression of chamber pressure against mass flow rate yields a slope c*/A_t = 700.00 kPa/(kg/s) corresponding to c* = 1750 m/s with R^2 > 0.999999.
Aeronautical apparatus & wind tunnel sensors.
High-Pressure Cryogenic Propellant Feed System
Pressurized run tanks supplying liquid oxygen and kerosene at pressures up to 10 MPa.
Coriolis Direct Mass Flow Meters
Measures oxidizer and fuel mass flow rates (m_dot_ox, m_dot_fuel) with 0.1% accuracy.
Flush-Mounted Fast Chamber Pressure Transducers
Measures static chamber pressure P_c and high-frequency combustion roughness.
Multi-Axis Thrust Vector Load Cell Frame
Measures axial thrust force F and side gimbal forces with 0.05% full-scale resolution.
Automated Engine FADEC Test Controller
Sequences cryogenic chilldown, spark torch ignition, and multi-step propellant throttling.
FOLLOW THE METHOD
A reason behind every step.
Measure propellant mass flow rate m_dot, chamber pressure P_c, and thrust F to determine characteristic exhaust velocity c* and thrust coefficient C_F. Standardized parameter settings: 1.5, 2.2, 3.0, 4.0, 5.0, 6.0 kg/s.
01
Cryogenic Chilldown and Purge
Execute gaseous nitrogen purge and liquid oxygen line chilldown until liquid temperature reaches 90 K at injector inlet.
02
Torch Igniter Activation
Ignite methane-oxygen torch pilot flame, verifying optical ignition sense before main propellant valve opening.
03
Main Valve Opening and Low-Throttle Fire
Open main LOX and RP-1 propellant valves to establish baseline firing at m_dot = 1.5 kg/s, stabilizing P_c = 1050 kPa.
04
Systematic Mass Flow Throttling
Throttle propellant flow through m_dot = 2.2, 3.0, 4.0, 5.0, and 6.0 kg/s while maintaining constant oxidizer-to-fuel ratio (O/F = 2.4).
05
Chamber Pressure and Thrust Recording
Record steady-state chamber pressure P_c and thrust F across 2000 samples per channel at each throttle setting.
06
Characteristic Velocity Linear Regression
Plot P_c (kPa) versus m_dot (kg/s), perform linear regression to determine slope c*/A_t, and compute c* from throat area A_t = 0.0025 m^2.
YOUR NOTEBOOK · 0 OBSERVATIONS
Turn measurements into aerodynamic insight.
Points are your recorded observations. The solid regression line validates Liquid Rocket Thrust Chamber mathematical scaling.
LINEAR REGRESSION FIT · y = mx + b
P_c = \left(\frac{c^*}{A_t}\right) \cdot \dot{m} \implies P_c = m \cdot \dot{m} + b
| # | Total Mass Flow Rate m_dot (kg/s) | Chamber Pressure P_c (kPa) | Total Propellant Mass Flow Rate m_dot (kg/s) | Combustion Chamber Pressure P_c (kPa) | Remove |
|---|
No readings yet. Execute an aerodynamic trial and record it.
Calculate it yourself.
CHECK YOUR UNDERSTANDING