[{"data":1,"prerenderedAt":-1},["ShallowReactive",2],{"skill:01M6CX0M3D7ZTF4RT2GFYJYNEC":3,"skill-same:01M6CX0M3D7ZTF4RT2GFYJYNEC":57},{"skillNo":4,"name":5,"nameZh":6,"description":7,"descriptionZh":6,"repositoryFullName":8,"ownerLogin":9,"ownerAvatarUrl":10,"ownerAccountNo":11,"ownerDisplayName":6,"repositoryNo":12,"starsCount":13,"licenseSpdxId":14,"licenseFileUrl":6,"sourceHtmlUrl":15,"categories":16,"keywords":17,"publishedAt":23,"linkOnly":24,"markdownContent":25,"markdownBody":26,"frontMatter":27,"refName":30,"commitSha":31,"skillDirectory":32,"forksCount":33,"repositoryHtmlUrl":34,"repositoryHomepageUrl":6,"ownerHtmlUrl":35,"ownerBlogUrl":6,"ownerDescription":6,"updatedAt":36,"install":37,"resources":50},"01M6CX0M3D7ZTF4RT2GFYJYNEC","aerospace-engineering",null,"Aerospace engineering fundamentals including aerodynamics, propulsion, flight dynamics, spacecraft dynamics, and structural analysis","ffsshhttiikk\u002Fopencode-agents-skills","ffsshhttiikk","https:\u002F\u002Favatars.githubusercontent.com\u002Fu\u002F226346827?v=4","01M6CWD9RR3Y7VRQX3MXYC8FYH","01M6CWD9WVPYAZW0J82G1VKGHY",3,"MIT","https:\u002F\u002Fgithub.com\u002Fffsshhttiikk\u002Fopencode-agents-skills\u002Fblob\u002Fmain\u002Fengineering\u002Faerospace-engineering\u002FSKILL.md",[],[18,19,20,21,22],"engineering","aerospace","aerospace engineering","analysis","opencode","2026-11-10 10:12:30",false,"---\nname: aerospace-engineering\ndescription: Aerospace engineering fundamentals including aerodynamics, propulsion, flight dynamics, spacecraft dynamics, and structural analysis\nlicense: MIT\ncompatibility: opencode\nmetadata:\n  audience: engineers\n  category: engineering\n---\n\n## What I do\n- Analyze aerodynamic forces and moments\n- Design aircraft and spacecraft propulsion systems\n- Model flight dynamics and stability\n- Calculate orbital mechanics and trajectories\n- Perform structural analysis for aerospace structures\n- Specify materials for aerospace applications\n- Analyze loads and fatigue for flight vehicles\n\n## When to use me\nWhen designing aircraft, spacecraft, propulsion systems, or analyzing aerodynamic and structural performance of aerospace vehicles.\n\n## Core Concepts\n- Aerodynamics and fluid dynamics\n- Propulsion systems (jet, rocket, hybrid)\n- Flight dynamics and control\n- Orbital mechanics and astrodynamics\n- Aerospace structures and materials\n- Load factors and fatigue analysis\n- Aircraft performance (range, endurance, climb)\n- Compressible flow and shock waves\n- Stability and control derivatives\n- Spacecraft attitude dynamics\n\n## Code Examples\n\n### Aerodynamic Calculations\n```python\nimport math\nfrom dataclasses import dataclass\n\n@dataclass\nclass Atmosphere:\n    altitude: float  # ft\n    temperature: float  # Rankine\n    pressure: float  # psf\n    density: float  # slugs\u002Fft³\n    speed_of_sound: float  # ft\u002Fs\n\ndef standard_atmosphere(altitude: float) -> Atmosphere:\n    \"\"\"Calculate standard atmosphere properties.\"\"\"\n    if altitude \u003C 36000:\n        T = 519.67 - 0.00356 * altitude\n        P = 2116.2 * (T \u002F 518.67)**5.256\n    else:\n        T = 389.97\n        P = 2116.2 * math.exp(-(altitude - 36000) \u002F 14700)\n    rho = P \u002F (1716 * T)\n    a = math.sqrt(1.4 * 1716 * T)\n    return Atmosphere(altitude, T, P, rho, a)\n\ndef lift_coefficient(\n    CL: float,\n    rho: float,\n    V: float,\n    S: float\n) -> float:\n    \"\"\"Calculate lift force.\"\"\"\n    return 0.5 * rho * V**2 * S * CL\n\ndef drag_polar(\n    CD0: float,\n    e: float,\n    AR: float,\n    CL: float\n) -> float:\n    \"\"\"Calculate drag coefficient using drag polar.\"\"\"\n    return CD0 + CL**2 \u002F (math.pi * e * AR)\n\ndef mach_number(\n    V: float,\n    a: float\n) -> float:\n    \"\"\"Calculate Mach number.\"\"\"\n    return V \u002F a\n\ndef dynamic_pressure(\n    rho: float,\n    V: float\n) -> float:\n    \"\"\"Calculate dynamic pressure.\"\"\"\n    return 0.5 * rho * V**2\n\n# Example: Aircraft performance at 20000 ft\nalt = 20000\natm = standard_atmosphere(alt)\nV = 400  # ktas\nCL = 0.8\nS = 500  # ft²\nq = dynamic_pressure(atm.density, V * 1.688)\nM = mach_number(V * 1.688, atm.speed_of_sound)\nprint(f\"Density: {atm.density:.6f} slugs\u002Fft³\")\nprint(f\"Dynamic pressure: {q:.1f} psf\")\nprint(f\"Mach number: {M:.3f}\")\n```\n\n### Flight Performance\n```python\ndef thrust_available(\n    TSL: float,  # sea level thrust\n    altitude: float,\n    velocity: float\n) -> float:\n    \"\"\"Calculate thrust at altitude using Napier's approximation.\"\"\"\n    sigma = standard_atmosphere(altitude).density \u002F 0.002377\n    return TSL * sigma * (1 - velocity**2 \u002F 5e10)\n\ndef power_available(\n    HP: float,  # shaft horsepower\n    prop_efficiency: float,\n    rho: float\n) -> float:\n    \"\"\"Calculate available power.\"\"\"\n    return HP * 550 * prop_efficiency \u002F (rho \u002F 0.002377)\n\ndef range_breguet(\n    R: float,  # range nm\n    V: float,  # true airspeed\n    L_D: float,  # lift-to-drag ratio\n    TSFC: float,  # thrust specific fuel consumption \u002Fhr\n    W_start: float,\n    W_end: float\n) -> float:\n    \"\"\"Calculate fuel required using Breguet range equation.\"\"\"\n    return V * L_D \u002F TSFC * math.log(W_start \u002F W_end)\n\ndef stall_speed(\n    W: float,  # weight lb\n    S: float,  # wing area ft²\n    CLmax: float,\n    rho: float\n) -> float:\n    \"\"\"Calculate stall speed.\"\"\"\n    Vs = math.sqrt(2 * W \u002F (rho * S * CLmax))\n    return Vs * 0.5925  # convert to knots\n\ndef climb_gradient(\n    T_W: float,  # thrust-to-weight ratio\n    W_S: float,  # wing loading lb\u002Fft²\n    rho: float,\n    CLmax: float\n) -> float:\n    \"\"\"Calculate maximum climb gradient.\"\"\"\n    return (T_W - W_S * 0.5 * rho * CLmax \u002F (2 * W_S)) \u002F (W_S * 0.5 \u002F (rho * CLmax))\n\n# Example: Aircraft range calculation\nW0 = 80000  # lb\nWf = 20000  # lb\nrange_nm = 3500\nL_D = 18\nV = 450  # kts\nTSFC = 0.5\nfuel_required = range_nm * TSFC \u002F (V * L_D) * (W0 - Wf)\nprint(f\"Fuel fraction required: {fuel_required:.4f}\")\n```\n\n### Rocket Propulsion\n```python\ndef rocket_thrust(\n    md: float,  # mass flow rate kg\u002Fs\n    Ve: float,  # exhaust velocity m\u002Fs\n    Ae: float,  # exit area m²\n    pe: float,  # exit pressure Pa\n    pa: float,  # ambient pressure Pa\n) -> float:\n    \"\"\"Calculate rocket thrust.\"\"\"\n    return md * Ve + (pe - pa) * Ae\n\ndef specific_impulse(\n    thrust: float,\n    md: float,\n    g0: float = 9.81\n) -> float:\n    \"\"\"Calculate specific impulse.\"\"\"\n    return thrust \u002F (md * g0)\n\ndef mass_ratio(\n    m_initial: float,\n    m_final: float\n) -> float:\n    \"\"\"Calculate mass ratio.\"\"\"\n    return m_initial \u002F m_final\n\ndef delta_v_rocket(\n    Isp: float,  # seconds\n    m_ratio: float,\n    g0: float = 9.81\n) -> float:\n    \"\"\"Calculate delta-V using rocket equation.\"\"\"\n    return Isp * g0 * math.log(m_ratio)\n\ndef nozzle_expansion_ratio(\n    pe: float,  # exit pressure Pa\n    pc: float,  # chamber pressure Pa\n    gamma: float = 1.2\n) -> float:\n    \"\"\"Calculate nozzle area expansion ratio.\"\"\"\n    return ((gamma + 1) \u002F 2)**((gamma + 1) \u002F (2 * (gamma - 1))) * math.sqrt(pe \u002F pc)\n\n# Example: Rocket calculation\nm_dot = 250  # kg\u002Fs\nVe = 3500  # m\u002Fs\nchamber_pressure = 7e6  # Pa\nIsp = 350\nm_ratio = 15\ndV = delta_v_rocket(Isp, m_ratio)\nprint(f\"Delta-V capability: {dV:.0f} m\u002Fs\")\n```\n\n### Orbital Mechanics\n```python\ndef orbital_velocity(\n    mu: float,  # gravitational parameter m³\u002Fs²\n    r: float  # orbital radius m\n) -> float:\n    \"\"\"Calculate circular orbital velocity.\"\"\"\n    return math.sqrt(mu \u002F r)\n\ndef orbital_period(\n    mu: float,\n    a: float  # semi-major axis m\n) -> float:\n    \"\"\"Calculate orbital period.\"\"\"\n    return 2 * math.pi * math.sqrt(a**3 \u002F mu)\n\ndef hohmann_transfer(\n    r1: float,  # initial radius m\n    r2: float,  # final radius m\n    mu: float\n) -> dict:\n    \"\"\"Calculate Hohmann transfer parameters.\"\"\"\n    a_transfer = (r1 + r2) \u002F 2\n    dv1 = math.sqrt(mu \u002F r1) * (math.sqrt(2 * r2 \u002F (r1 + r2)) - 1)\n    dv2 = math.sqrt(mu \u002F r2) * (1 - math.sqrt(2 * r1 \u002F (r1 + r2)))\n    transfer_time = math.pi * math.sqrt(a_transfer**3 \u002F mu)\n    return {\n        \"dv1\": dv1,\n        \"dv2\": dv2,\n        \"total_dv\": dv1 + dv2,\n        \"transfer_time\": transfer_time\n    }\n\ndef escape_velocity(\n    mu: float,\n    r: float\n) -> float:\n    \"\"\"Calculate escape velocity.\"\"\"\n    return math.sqrt(2 * mu \u002F r)\n\n# Example: GEO transfer\nr_leo = 6678e3  # m\nr_geo = 42164e3  # m\nmu_earth = 3.986e14  # m³\u002Fs²\ntransfer = hohmann_transfer(r_leo, r_geo, mu_earth)\nprint(f\"Transfer burn: {transfer['dv1']:.0f} m\u002Fs\")\nprint(f\"Insertion burn: {transfer['dv2']:.0f} m\u002Fs\")\nprint(f\"Total delta-V: {transfer['total_dv']:.0f} m\u002Fs\")\n```\n\n## Best Practices\n- Use consistent unit systems throughout calculations\n- Apply appropriate factors of safety for flight hardware\n- Consider environmental factors (temperature, pressure) in performance\n- Validate analytical results with CFD and wind tunnel data\n- Follow aerospace standards (MIL, NASA, FAA, EASA)\n- Account for structural loads in aerodynamic design\n- Consider maintainability and inspection requirements\n- Use margin on critical performance parameters\n- Consider failure modes and safety margins\n- Document all assumptions and methods used\n","## What I do\n- Analyze aerodynamic forces and moments\n- Design aircraft and spacecraft propulsion systems\n- Model flight dynamics and stability\n- Calculate orbital mechanics and trajectories\n- Perform structural analysis for aerospace structures\n- Specify materials for aerospace applications\n- Analyze loads and fatigue for flight vehicles\n\n## When to use me\nWhen designing aircraft, spacecraft, propulsion systems, or analyzing aerodynamic and structural performance of aerospace vehicles.\n\n## Core Concepts\n- Aerodynamics and fluid dynamics\n- Propulsion systems (jet, rocket, hybrid)\n- Flight dynamics and control\n- Orbital mechanics and astrodynamics\n- Aerospace structures and materials\n- Load factors and fatigue analysis\n- Aircraft performance (range, endurance, climb)\n- Compressible flow and shock waves\n- Stability and control derivatives\n- Spacecraft attitude dynamics\n\n## Code Examples\n\n### Aerodynamic Calculations\n```python\nimport math\nfrom dataclasses import dataclass\n\n@dataclass\nclass Atmosphere:\n    altitude: float  # ft\n    temperature: float  # Rankine\n    pressure: float  # psf\n    density: float  # slugs\u002Fft³\n    speed_of_sound: float  # ft\u002Fs\n\ndef standard_atmosphere(altitude: float) -> Atmosphere:\n    \"\"\"Calculate standard atmosphere properties.\"\"\"\n    if altitude \u003C 36000:\n        T = 519.67 - 0.00356 * altitude\n        P = 2116.2 * (T \u002F 518.67)**5.256\n    else:\n        T = 389.97\n        P = 2116.2 * math.exp(-(altitude - 36000) \u002F 14700)\n    rho = P \u002F (1716 * T)\n    a = math.sqrt(1.4 * 1716 * T)\n    return Atmosphere(altitude, T, P, rho, a)\n\ndef lift_coefficient(\n    CL: float,\n    rho: float,\n    V: float,\n    S: float\n) -> float:\n    \"\"\"Calculate lift force.\"\"\"\n    return 0.5 * rho * V**2 * S * CL\n\ndef drag_polar(\n    CD0: float,\n    e: float,\n    AR: float,\n    CL: float\n) -> float:\n    \"\"\"Calculate drag coefficient using drag polar.\"\"\"\n    return CD0 + CL**2 \u002F (math.pi * e * AR)\n\ndef mach_number(\n    V: float,\n    a: float\n) -> float:\n    \"\"\"Calculate Mach number.\"\"\"\n    return V \u002F a\n\ndef dynamic_pressure(\n    rho: float,\n    V: float\n) -> float:\n    \"\"\"Calculate dynamic pressure.\"\"\"\n    return 0.5 * rho * V**2\n\n# Example: Aircraft performance at 20000 ft\nalt = 20000\natm = standard_atmosphere(alt)\nV = 400  # ktas\nCL = 0.8\nS = 500  # ft²\nq = dynamic_pressure(atm.density, V * 1.688)\nM = mach_number(V * 1.688, atm.speed_of_sound)\nprint(f\"Density: {atm.density:.6f} slugs\u002Fft³\")\nprint(f\"Dynamic pressure: {q:.1f} psf\")\nprint(f\"Mach number: {M:.3f}\")\n```\n\n### Flight Performance\n```python\ndef thrust_available(\n    TSL: float,  # sea level thrust\n    altitude: float,\n    velocity: float\n) -> float:\n    \"\"\"Calculate thrust at altitude using Napier's approximation.\"\"\"\n    sigma = standard_atmosphere(altitude).density \u002F 0.002377\n    return TSL * sigma * (1 - velocity**2 \u002F 5e10)\n\ndef power_available(\n    HP: float,  # shaft horsepower\n    prop_efficiency: float,\n    rho: float\n) -> float:\n    \"\"\"Calculate available power.\"\"\"\n    return HP * 550 * prop_efficiency \u002F (rho \u002F 0.002377)\n\ndef range_breguet(\n    R: float,  # range nm\n    V: float,  # true airspeed\n    L_D: float,  # lift-to-drag ratio\n    TSFC: float,  # thrust specific fuel consumption \u002Fhr\n    W_start: float,\n    W_end: float\n) -> float:\n    \"\"\"Calculate fuel required using Breguet range equation.\"\"\"\n    return V * L_D \u002F TSFC * math.log(W_start \u002F W_end)\n\ndef stall_speed(\n    W: float,  # weight lb\n    S: float,  # wing area ft²\n    CLmax: float,\n    rho: float\n) -> float:\n    \"\"\"Calculate stall speed.\"\"\"\n    Vs = math.sqrt(2 * W \u002F (rho * S * CLmax))\n    return Vs * 0.5925  # convert to knots\n\ndef climb_gradient(\n    T_W: float,  # thrust-to-weight ratio\n    W_S: float,  # wing loading lb\u002Fft²\n    rho: float,\n    CLmax: float\n) -> float:\n    \"\"\"Calculate maximum climb gradient.\"\"\"\n    return (T_W - W_S * 0.5 * rho * CLmax \u002F (2 * W_S)) \u002F (W_S * 0.5 \u002F (rho * CLmax))\n\n# Example: Aircraft range calculation\nW0 = 80000  # lb\nWf = 20000  # lb\nrange_nm = 3500\nL_D = 18\nV = 450  # kts\nTSFC = 0.5\nfuel_required = range_nm * TSFC \u002F (V * L_D) * (W0 - Wf)\nprint(f\"Fuel fraction required: {fuel_required:.4f}\")\n```\n\n### Rocket Propulsion\n```python\ndef rocket_thrust(\n    md: float,  # mass flow rate kg\u002Fs\n    Ve: float,  # exhaust velocity m\u002Fs\n    Ae: float,  # exit area m²\n    pe: float,  # exit pressure Pa\n    pa: float,  # ambient pressure Pa\n) -> float:\n    \"\"\"Calculate rocket thrust.\"\"\"\n    return md * Ve + (pe - pa) * Ae\n\ndef specific_impulse(\n    thrust: float,\n    md: float,\n    g0: float = 9.81\n) -> float:\n    \"\"\"Calculate specific impulse.\"\"\"\n    return thrust \u002F (md * g0)\n\ndef mass_ratio(\n    m_initial: float,\n    m_final: float\n) -> float:\n    \"\"\"Calculate mass ratio.\"\"\"\n    return m_initial \u002F m_final\n\ndef delta_v_rocket(\n    Isp: float,  # seconds\n    m_ratio: float,\n    g0: float = 9.81\n) -> float:\n    \"\"\"Calculate delta-V using rocket equation.\"\"\"\n    return Isp * g0 * math.log(m_ratio)\n\ndef nozzle_expansion_ratio(\n    pe: float,  # exit pressure Pa\n    pc: float,  # chamber pressure Pa\n    gamma: float = 1.2\n) -> float:\n    \"\"\"Calculate nozzle area expansion ratio.\"\"\"\n    return ((gamma + 1) \u002F 2)**((gamma + 1) \u002F (2 * (gamma - 1))) * math.sqrt(pe \u002F pc)\n\n# Example: Rocket calculation\nm_dot = 250  # kg\u002Fs\nVe = 3500  # m\u002Fs\nchamber_pressure = 7e6  # Pa\nIsp = 350\nm_ratio = 15\ndV = delta_v_rocket(Isp, m_ratio)\nprint(f\"Delta-V capability: {dV:.0f} m\u002Fs\")\n```\n\n### Orbital Mechanics\n```python\ndef orbital_velocity(\n    mu: float,  # gravitational parameter m³\u002Fs²\n    r: float  # orbital radius m\n) -> float:\n    \"\"\"Calculate circular orbital velocity.\"\"\"\n    return math.sqrt(mu \u002F r)\n\ndef orbital_period(\n    mu: float,\n    a: float  # semi-major axis m\n) -> float:\n    \"\"\"Calculate orbital period.\"\"\"\n    return 2 * math.pi * math.sqrt(a**3 \u002F mu)\n\ndef hohmann_transfer(\n    r1: float,  # initial radius m\n    r2: float,  # final radius m\n    mu: float\n) -> dict:\n    \"\"\"Calculate Hohmann transfer parameters.\"\"\"\n    a_transfer = (r1 + r2) \u002F 2\n    dv1 = math.sqrt(mu \u002F r1) * (math.sqrt(2 * r2 \u002F (r1 + r2)) - 1)\n    dv2 = math.sqrt(mu \u002F r2) * (1 - math.sqrt(2 * r1 \u002F (r1 + r2)))\n    transfer_time = math.pi * math.sqrt(a_transfer**3 \u002F mu)\n    return {\n        \"dv1\": dv1,\n        \"dv2\": dv2,\n        \"total_dv\": dv1 + dv2,\n        \"transfer_time\": transfer_time\n    }\n\ndef escape_velocity(\n    mu: float,\n    r: float\n) -> float:\n    \"\"\"Calculate escape velocity.\"\"\"\n    return math.sqrt(2 * mu \u002F r)\n\n# Example: GEO transfer\nr_leo = 6678e3  # m\nr_geo = 42164e3  # m\nmu_earth = 3.986e14  # m³\u002Fs²\ntransfer = hohmann_transfer(r_leo, r_geo, mu_earth)\nprint(f\"Transfer burn: {transfer['dv1']:.0f} m\u002Fs\")\nprint(f\"Insertion burn: {transfer['dv2']:.0f} m\u002Fs\")\nprint(f\"Total delta-V: {transfer['total_dv']:.0f} m\u002Fs\")\n```\n\n## Best Practices\n- Use consistent unit systems throughout calculations\n- Apply appropriate factors of safety for flight hardware\n- Consider environmental factors (temperature, pressure) in performance\n- Validate analytical results with CFD and wind tunnel data\n- Follow aerospace standards (MIL, NASA, FAA, EASA)\n- Account for structural loads in aerodynamic design\n- Consider maintainability and inspection requirements\n- Use margin on critical performance parameters\n- Consider failure modes and safety margins\n- Document all assumptions and methods 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