Model Rocket Flight Simulator & Stability Calculator

Powered flight solves $\dot h=v$, $\dot v=(T-kv|v|)/m-g$ and $\dot J=T$ with adaptive Dormand–Prince integration. Here $k=\rho C_D A/2$ and $m=m_d+m_p(1-J/I)$: consumed propellant is proportional to cumulative impulse. Motor samples are linearly interpolated and trapezoidal integration gives total impulse. Integration restarts at every thrust-curve knot, carrying altitude, velocity and accumulated impulse across each boundary, so narrow thrust spikes cannot be skipped. The current-input impulse check compares numerical consumption with the exact integral of the linear segments; it blocks results above a relative discrepancy of 0.0001.

Liftoff is found from thrust minus the changing weight while the rocket remains on the pad. After that instant, pad holding is disabled. The first downward ground crossing is refined by bisection, the recorded trajectory ends at zero height, and further thrust cannot relaunch it. Ground impact during the burn and upward velocity reversing below −0.00001 m/s before burnout are rejected. The coast/apogee and ejection formulas apply only to a continuous ascent through burnout.

After burnout, constant mass and density permit exact quadratic-drag coast equations. For $q=\sqrt{m_d g/k}$, the ideal delay is $(q/g)\tan^{-1}(v_b/q)$ and free-flight apogee adds $m_d\ln(1+kv_b^2/(m_dg))/(2k)$ to burnout altitude. The zero-drag limit uses ballistic equations. Selected ejection may precede or follow free-flight apogee. The recovery solution switches instantly to canopy drag, handles residual upward motion and solves for touchdown without a fixed terminal-speed approximation.

Stability uses the simplified low-speed Barrowman nose-plus-fin method: nose normal-force factor 2; nose CP at two-thirds nose length for a cone or 0.466 for a tangent ogive; finite-span trapezoidal fin factor with fin–body interference. Component moments are divided by total normal-force factor to locate CP. Uniform body lift and transitions are omitted. NASA TM-2001-209983, attached pp. 12 and 16, gives the small-angle assumptions and quarter-chord basis. CG at burnout removes the propellant's mass moment. Static margin is $(x_{CP}-x_{CG})/D$; its minimum occurs at one of the two burn endpoints under the assumed fixed propellant centroid.

Canopy sizing follows force balance $m_dg=\rho C_{D,ch}A_{ch}v_t^2/2$, with $A_{ch}=\pi D_{ch}^2/4$. The recovery solution and wind drift are derived here from that balance and Newton's law.

• One stage; thrust remains vertical. No wind force or attitude dynamics during ascent.
• Constant air density, airframe drag coefficient and standard Earth gravity (9.80665 m/s²). A fixed 340.3 m/s sound speed is used only to screen the low-Mach range.
• Burnout mass includes the entire recovering vehicle and empty casing. Ejection-charge mass loss is omitted.
• Propellant consumption follows cumulative impulse, with a fixed propellant centroid. Airframe and spent motor mass distribution remain fixed.
• Three or four identical thin trapezoidal fins, a straight circular body and the selected ideal nose. The measured CG is independent of the drawing.
• The canopy opens completely and instantaneously. Its drag dominates airframe drag after deployment. Its reference diameter must match its quoted Cd; 0.75 is a flat-parasheet estimate.
• Wind drift equals wind speed multiplied by time after deployment; no ascent drift.
• The default and idealized motor shapes are original mathematical profiles, not manufacturer data.

For preliminary comparisons of slender, rigid, single-stage model rockets at small angle of attack, below roughly Mach 0.5 and 1 km altitude. The model requires continuous ascent after liftoff through burnout and deployment before ground impact. Any descent during the burn beyond a 0.00001 m/s numerical velocity tolerance is unsupported, even when later thrust restores upward motion. A downward ground crossing during the burn terminates the trajectory and blocks the flight predictions. The one-caliber margin and 20 m/s deployment-speed warning are screening cues, not certified limits.

Six manually entered interior samples can underresolve a complex .eng curve. Check integrated impulse and curve shape against your original. RASP header motor masses are not the airframe mass. No direct .eng uploader is available in this viewer.

No transonic aerodynamics, multiple fin sets, body transitions, boat tails, supersonic flight, weathercocking, rail-exit stability, canopy inflation, opening shock, structural loads or launch-site clearance prediction. A CP calculation alone cannot prove dynamic stability. Compare with a detailed flight simulator and measured flights before selecting a real motor delay. Follow local launch rules and the motor and recovery manufacturers' instructions.