Aerospace MCP

Instructions

Implementation Reference

• aerospace_mcp/tools/rockets.py:10-109 (handler)

  MCP tool handler function that processes input parameters (rocket_geometry dict, launch_conditions dict, simulation_options dict), creates RocketGeometry object, delegates to core trajectory analysis, and formats results into readable text with performance summary, key trajectory points, and JSON performance data.
  def rocket_3dof_trajectory( rocket_geometry: dict, launch_conditions: dict, simulation_options: dict | None = None, ) -> str: """Calculate 3DOF rocket trajectory using numerical integration. Args: rocket_geometry: Rocket geometry parameters launch_conditions: Launch conditions (launch_angle_deg, launch_site, etc.) simulation_options: Optional simulation settings Returns: Formatted string with trajectory analysis results """ try: from ..integrations.rockets import ( RocketGeometry, ) from ..integrations.rockets import ( rocket_3dof_trajectory as _trajectory_analysis, ) # Create geometry object geometry = RocketGeometry(**rocket_geometry) # Run trajectory analysis result = _trajectory_analysis(geometry, launch_conditions, simulation_options) # Format response result_lines = [ "3DOF Rocket Trajectory Analysis", "=" * 50, f"Rocket: {geometry.length_m:.1f}m long, {geometry.diameter_m:.2f}m dia", f"Mass: {geometry.dry_mass_kg:.1f} kg dry + {geometry.propellant_mass_kg:.1f} kg prop", f"Launch Angle: {launch_conditions.get('launch_angle_deg', 90):.1f}°", "", "Performance Summary:", f" Max Altitude: {result.performance.max_altitude_m:.0f} m ({result.performance.max_altitude_m / 1000:.2f} km)", f" Max Velocity: {result.performance.max_velocity_ms:.1f} m/s (Mach {result.performance.max_mach:.2f})", f" Apogee Time: {result.performance.apogee_time_s:.1f} s", f" Burnout Time: {result.performance.burnout_time_s:.1f} s", f" Max Q: {result.performance.max_q_pa / 1000:.1f} kPa", f" Total Impulse: {result.performance.total_impulse_ns:.0f} N⋅s", f" Specific Impulse: {result.performance.specific_impulse_s:.1f} s", ] # Add trajectory points summary if hasattr(result, "trajectory") and result.trajectory: result_lines.extend( [ "", "Key Trajectory Points:", f"{'Time (s)':>8} {'Alt (m)':>8} {'Vel (m/s)':>10} {'Accel (m/s²)':>12}", ] ) result_lines.append("-" * 45) # Sample key points trajectory = result.trajectory sample_indices = [ 0, len(trajectory) // 4, len(trajectory) // 2, 3 * len(trajectory) // 4, -1, ] for i in sample_indices: if i < len(trajectory): point = trajectory[i] result_lines.append( f"{point.time_s:8.1f} {point.altitude_m:8.0f} {point.velocity_ms:10.1f} " f"{point.acceleration_ms2:12.1f}" ) # Add JSON data json_data = json.dumps( { "performance": { "max_altitude_m": result.performance.max_altitude_m, "max_velocity_ms": result.performance.max_velocity_ms, "apogee_time_s": result.performance.apogee_time_s, "burnout_time_s": result.performance.burnout_time_s, "max_q_pa": result.performance.max_q_pa, "total_impulse_ns": result.performance.total_impulse_ns, "specific_impulse_s": result.performance.specific_impulse_s, } }, indent=2, ) result_lines.extend(["", "JSON Performance Data:", json_data]) return "\n".join(result_lines) except ImportError: return "Rocket trajectory analysis not available - install rocketry packages" except Exception as e: logger.error(f"Rocket trajectory error: {str(e)}", exc_info=True) return f"Rocket trajectory error: {str(e)}"
• aerospace_mcp/tools/agents.py:154-166 (schema)

  Explicit schema definition for the rocket_3dof_trajectory tool as a ToolReference, including name, description, parameters (geometry dict, dt_s, max_time_s, launch_angle_deg), and examples.
  ToolReference( name="rocket_3dof_trajectory", description="Simulate 3DOF rocket trajectory with atmospheric effects", parameters={ "geometry": "dict - Rocket geometry with mass_kg, thrust_n, burn_time_s, drag_coeff, reference_area_m2", "dt_s": "float - Time step in seconds (default 0.1)", "max_time_s": "float - Maximum simulation time in seconds (default 300)", "launch_angle_deg": "float - Launch angle in degrees from vertical (default 0)", }, examples=[ 'rocket_3dof_trajectory({"mass_kg": 500, "thrust_n": 8000, "burn_time_s": 60, "drag_coeff": 0.3, "reference_area_m2": 0.5}, 0.1, 300, 15)' ], ),
• aerospace_mcp/fastmcp_server.py:111-111 (registration)

  Explicit registration of the rocket_3dof_trajectory tool with FastMCP server via mcp.tool(rocket_3dof_trajectory).
  mcp.tool(rocket_3dof_trajectory)
• aerospace_mcp/integrations/rockets.py:94-228 (helper)

  Core physics implementation: 3DOF rocket trajectory simulator using numerical Euler integration, handling thrust curves, atmospheric density, drag forces, gravity, propellant mass flow, and returning list of trajectory points.
  def rocket_3dof_trajectory( geometry: RocketGeometry, dt_s: float = 0.1, max_time_s: float = 300.0, launch_angle_deg: float = 90.0, ) -> list[RocketTrajectoryPoint]: """ Calculate 3DOF rocket trajectory using numerical integration. Args: geometry: Rocket geometry and mass properties dt_s: Time step for integration (seconds) max_time_s: Maximum simulation time (seconds) launch_angle_deg: Launch angle from horizontal (degrees) Returns: List of trajectory points """ # Initial conditions trajectory = [] time = 0.0 altitude = 0.0 # m above sea level velocity_vertical = 0.0 # m/s velocity_horizontal = 0.0 # m/s mass = geometry.dry_mass_kg + geometry.propellant_mass_kg # Launch angle in radians launch_angle_rad = math.radians(launch_angle_deg) # Constants g0 = 9.80665 # m/s² standard gravity # Get atmosphere profile for drag calculations max_alt_m = 50000 # 50km max altitude for atmosphere alt_points = list(range(0, max_alt_m + 1000, 1000)) atm_profile = get_atmosphere_profile(alt_points, "ISA") atm_dict = {point.altitude_m: point for point in atm_profile} while time <= max_time_s and altitude >= 0: # Get current thrust thrust = ( get_thrust_at_time(geometry.thrust_curve, time) if geometry.thrust_curve else 0.0 ) # Get atmospheric properties at current altitude atm_alt = min(max_alt_m, max(0, int(altitude // 1000) * 1000)) if atm_alt in atm_dict: atm = atm_dict[atm_alt] rho = atm.density_kg_m3 speed_of_sound = atm.speed_of_sound_mps else: # Fallback for extreme altitudes rho = 1.225 * math.exp(-altitude / 8400) # Simple exponential model speed_of_sound = 343.0 * math.sqrt( max(0.1, (288.15 - 0.0065 * altitude) / 288.15) ) # Total velocity velocity_total = math.sqrt(velocity_vertical**2 + velocity_horizontal**2) # Mach number mach = velocity_total / speed_of_sound if speed_of_sound > 0 else 0.0 # Dynamic pressure dynamic_pressure = 0.5 * rho * velocity_total**2 # Drag force (opposing velocity direction) drag_area = math.pi * (geometry.diameter_m / 2) ** 2 drag_force = geometry.cd * drag_area * dynamic_pressure # Drag acceleration components if velocity_total > 0: drag_accel_vertical = ( -drag_force * (velocity_vertical / velocity_total) / mass ) drag_accel_horizontal = ( -drag_force * (velocity_horizontal / velocity_total) / mass ) else: drag_accel_vertical = 0.0 drag_accel_horizontal = 0.0 # Thrust acceleration (in launch direction initially, then vertical) if time < 5.0: # First 5 seconds follow launch angle thrust_accel_vertical = thrust * math.sin(launch_angle_rad) / mass thrust_accel_horizontal = thrust * math.cos(launch_angle_rad) / mass else: # After 5 seconds, thrust is vertical only thrust_accel_vertical = thrust / mass thrust_accel_horizontal = 0.0 # Total acceleration accel_vertical = thrust_accel_vertical - g0 + drag_accel_vertical accel_horizontal = thrust_accel_horizontal + drag_accel_horizontal # Record current state trajectory.append( RocketTrajectoryPoint( time_s=time, altitude_m=altitude, velocity_ms=velocity_total, acceleration_ms2=math.sqrt(accel_vertical**2 + accel_horizontal**2), mass_kg=mass, thrust_n=thrust, drag_n=drag_force, mach=mach, dynamic_pressure_pa=dynamic_pressure, ) ) # Integration (Euler method) velocity_vertical += accel_vertical * dt_s velocity_horizontal += accel_horizontal * dt_s altitude += velocity_vertical * dt_s # Update mass (simple propellant consumption) if thrust > 0 and geometry.thrust_curve: # Estimate mass flow rate from thrust curve # For a simplified model, assume constant mass flow rate during burn burn_time_total = max( (point[0] for point in geometry.thrust_curve if point[1] > 0), default=1.0, ) if burn_time_total > 0: mass_flow_rate = geometry.propellant_mass_kg / burn_time_total mass = max(geometry.dry_mass_kg, mass - mass_flow_rate * dt_s) time += dt_s # Stop if apogee reached and descending if altitude > 0 and velocity_vertical < -1.0: break return trajectory
• aerospace_mcp/integrations/rockets.py:15-27 (helper)

  Dataclass RocketGeometry defining structure for rocket parameters: dry_mass_kg, propellant_mass_kg, diameter_m, length_m, cd, thrust_curve.
  @dataclass class RocketGeometry: """Rocket geometry parameters.""" dry_mass_kg: float # Rocket dry mass propellant_mass_kg: float # Initial propellant mass diameter_m: float # Rocket diameter length_m: float # Total rocket length cd: float = 0.3 # Drag coefficient thrust_curve: list[list[float]] = ( None # [[time_s, thrust_N], ...] or constant thrust )