Heavy-Wall Ballistic Test Rocket Motor (BATES)
Re-usable stainless steel test motor with interchangeable converging-diverging graphite nozzles.
Aerospace & Aeronautical Engineering / Solid Rocket Internal Ballistics
FREE TO EXPLORELABORATORY 32 · AEROSPACE & AERONAUTICAL ENGINEERING
Solid Rocket Motor Internal Ballistics and Propellant Grain Geometry
Measure chamber pressure P_c(t) and thrust F(t) as a function of propellant burning surface area A_b(t) and Saint-Robert's burn rate law r = a * P_c^n. Explore interactive wind tunnel simulation, flight dynamics, and aerospace engineering telemetry.
AERODYNAMICS · PROPULSION · FLIGHT DYNAMICS
Your virtual aerospace engineering laboratory.
LINEAR BURNING RATE R—
FLIGHT & TEST CELL TELEMETRY
Ready to sample with P_c = 5.0 MPa.
Adjust P_c, 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
Erosive burning occurs near the nozzle entrance where high-velocity gas scouring increases local burn rate above Saint-Robert's law. Linear regression of ln(r) against ln(P_c) yields a pressure exponent slope n = 0.3500 with R^2 > 0.999999.
Aeronautical apparatus & wind tunnel sensors.
Hydroxyl-Terminated Polybutadiene (HTPB) Propellant
Standard AP/HTPB composite propellant grain with calibrated burning rate inhibitors.
High-Pressure Piezoelectric Chamber Transducer
Measures internal motor pressure P_c up to 20 MPa with 100 kHz sampling frequency.
Static Thrust Load Cell Stand
Measures axial reaction thrust F(t) with 0.1% accuracy and mechanical dampers.
Pyrotechnic Ignition System & Sequencer
Delivers uniform boron-potassium nitrate (BPN) squib ignition pulse across internal grain core.
Ballistics Data Acquisition Workstation
Executes pressure-time integration to compute total impulse I_tot and burning rate r.
FOLLOW THE METHOD
A reason behind every step.
Measure chamber pressure P_c(t) and thrust F(t) as a function of propellant burning surface area A_b(t) and Saint-Robert's burn rate law r = a * P_c^n. Standardized parameter settings: 2.5, 3.5, 5.0, 6.5, 8.0, 10.0 MPa.
01
Motor Assembly and Throat Measurement
Measure graphite nozzle throat diameter with optical micrometer (A_t = 125 mm^2) and assemble BATES grain.
02
Safety Interlock and Bunker Evacuation
Arm pyrotechnic squib circuit and evacuate personnel to blast bunker, verifying sensor calibration.
03
Static Firing Execution
Ignite the motor, sampling chamber pressure P_c(t) and thrust F(t) at 10,000 samples per second.
04
Action Time and Burn Rate Extraction
Determine burn action time t_b from the 75% pressure-time inflection points, calculating average burn rate r = w_web / t_b.
05
Multi-Pressure Throat Variation Sweep
Repeat tests across modular nozzle throats to achieve steady chamber pressures P_c = 2.5, 3.5, 5.0, 6.5, 8.0, and 10.0 MPa.
06
Saint-Robert Log-Log Linear Regression
Plot ln(r) versus ln(P_c), perform linear regression to extract pressure exponent n as the slope, and verify combustion stability (n < 1.0).
YOUR NOTEBOOK · 0 OBSERVATIONS
Turn measurements into aerodynamic insight.
Points are your recorded observations. The solid regression line validates Solid Rocket Internal Ballistics mathematical scaling.
LINEAR REGRESSION FIT · y = mx + b
\ln(r) = \ln(a) + n \ln(P_c) \implies \ln(r) = m \cdot \ln(P_c) + b
| # | Chamber Pressure P_c (MPa) | Linear Burning Rate r (mm/s) | Logarithm of Chamber Pressure ln(P_c [MPa]) | Logarithm of Propellant Burning Rate ln(r [mm/s]) | Remove |
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