La entrada atmosférica es el movimiento de un objeto desde el espacio exterior hacia y a través de los gases de la atmósfera de un planeta , planeta enano o satélite natural . Hay dos tipos principales de entrada atmosférica: entrada incontrolada, como la entrada de objetos astronómicos , desechos espaciales o bólidos ; y entrada (o reentrada) controlada de una nave espacial capaz de navegar o seguir un rumbo predeterminado. Las tecnologías y procedimientos que permiten la entrada, el descenso y el aterrizaje atmosféricos controlados de naves espaciales se denominan colectivamente EDL .
Los objetos que entran en una atmósfera experimentan un arrastre atmosférico , que ejerce presión mecánica sobre el objeto y un calentamiento aerodinámico, causado principalmente por la compresión del aire frente al objeto, pero también por el arrastre. Estas fuerzas pueden causar pérdida de masa ( ablación ) o incluso la desintegración completa de objetos más pequeños, y los objetos con menor resistencia a la compresión pueden explotar.
Los vehículos espaciales con tripulación deben reducir la velocidad a velocidades subsónicas antes de que se puedan desplegar los paracaídas o los frenos de aire. Dichos vehículos tienen energías cinéticas típicamente entre 50 y 1800 megajulios, y la disipación atmosférica es la única forma de gastar la energía cinética. La cantidad de combustible para cohetes necesaria para reducir la velocidad del vehículo sería casi igual a la cantidad utilizada para acelerarlo inicialmente, por lo que es muy poco práctico utilizar retrocohetes para todo el procedimiento de reentrada a la Tierra. Si bien la alta temperatura generada en la superficie del escudo térmico se debe a la compresión adiabática , la energía cinética del vehículo se pierde finalmente por la fricción del gas (viscosidad) después de que el vehículo ha pasado. Otras pérdidas de energía más pequeñas incluyen la radiación de cuerpo negro directamente de los gases calientes y las reacciones químicas entre los gases ionizados.
Las ojivas balísticas y los vehículos prescindibles no necesitan reducir la velocidad al volver a entrar y, de hecho, se hacen aerodinámicos para mantener su velocidad. Además, los retornos de baja velocidad a la Tierra desde el espacio cercano, como los saltos en paracaídas desde globos , no requieren protección térmica porque la aceleración gravitacional de un objeto que comienza en reposo relativo desde dentro de la atmósfera misma (o no muy por encima de ella) no puede generar suficiente velocidad. para causar un calentamiento atmosférico significativo.
Para la Tierra, la entrada atmosférica ocurre por convención en la línea Kármán a una altitud de 100 km (62 millas; 54 millas náuticas) sobre la superficie, mientras que en Venus la entrada atmosférica ocurre a 250 km (160 millas; 130 millas náuticas) y en Marte la entrada atmosférica entrada a unos 80 km (50 millas; 43 millas náuticas). Los objetos incontrolados alcanzan altas velocidades mientras aceleran a través del espacio hacia la Tierra bajo la influencia de la gravedad de la Tierra , y se ralentizan por la fricción al encontrar la atmósfera terrestre. Los meteoritos también viajan a menudo bastante rápido en relación con la Tierra simplemente porque su propia trayectoria orbital es diferente a la de la Tierra antes de encontrar el pozo de gravedad de la Tierra . La mayoría de los objetos controlados entran a velocidades hipersónicas debido a sus trayectorias suborbitales (por ejemplo, vehículos de reentrada de misiles balísticos intercontinentales ), orbitales (por ejemplo, Soyuz ) o ilimitadas (por ejemplo, meteoritos ). Se han desarrollado varias tecnologías avanzadas para permitir la reentrada atmosférica y el vuelo a velocidades extremas. Un método alternativo de entrada atmosférica controlada a baja velocidad es la flotabilidad [1], que es adecuada para la entrada planetaria donde atmósferas espesas, gravedad fuerte o ambos factores complican la entrada hiperbólica de alta velocidad, como las atmósferas de Venus , Titán y los gigantes gaseosos. . [2]
Historia
El concepto de escudo térmico ablativo fue descrito ya en 1920 por Robert Goddard : "En el caso de los meteoros, que entran en la atmósfera con velocidades de hasta 48 km (30 millas) por segundo, el interior de los meteoros permanece frío, y la erosión se debe, en gran medida, al desconchado o agrietamiento de la superficie repentinamente calentada. Por esta razón, si la superficie exterior del aparato estuviera formada por capas de una sustancia dura muy infusible con capas de un mal conductor de calor en el medio, la superficie no se erosionaría en un grado considerable, especialmente porque la velocidad del aparato no sería tan grande como la de un meteoro promedio ". [3]
El desarrollo práctico de los sistemas de reentrada comenzó a medida que aumentaba el alcance y la velocidad de reentrada de los misiles balísticos . Para los primeros misiles de corto alcance, como el V-2 , la estabilización y el estrés aerodinámico eran cuestiones importantes (muchos V-2 se rompían durante la reentrada), pero el calentamiento no era un problema grave. Los misiles de mediano alcance como el R-5 soviético , con un alcance de 1.200 kilómetros (650 millas náuticas), requerían protección térmica compuesta de cerámica en vehículos de reentrada separables (ya no era posible que toda la estructura del cohete sobreviviera al reentrada). Los primeros misiles balísticos intercontinentales , con alcances de 8.000 a 12.000 km (4.300 a 6.500 nmi), solo fueron posibles con el desarrollo de modernos escudos térmicos ablativos y vehículos de forma roma.
En los Estados Unidos, esta tecnología fue iniciado por el H. Julian Allen y AJ Eggers Jr. del Comité Consultivo Nacional de Aeronáutica (NACA) en el Centro de Investigación Ames . [4] En 1951, hicieron el descubrimiento contradictorio de que una forma roma (alta resistencia) era el escudo térmico más eficaz. [5] A partir de principios de ingeniería simples, Allen y Eggers demostraron que la carga de calor experimentada por un vehículo de entrada era inversamente proporcional al coeficiente de resistencia ; es decir, cuanto mayor es la resistencia, menor es la carga térmica. Si el vehículo de reentrada se embota, el aire no puede "salir del camino" lo suficientemente rápido y actúa como un colchón de aire para empujar la onda de choque y la capa de choque calentado hacia adelante (lejos del vehículo). Dado que la mayoría de los gases calientes ya no están en contacto directo con el vehículo, la energía térmica permanecería en el gas impactado y simplemente se movería alrededor del vehículo para luego disiparse en la atmósfera.
El descubrimiento de Allen y Eggers, aunque inicialmente tratado como un secreto militar, fue finalmente publicado en 1958. [6]
Terminología, definiciones y jerga
Durante las décadas transcurridas desde la década de 1950, ha crecido una rica jerga técnica en torno a la ingeniería de vehículos diseñados para entrar en atmósferas planetarias. Se recomienda que el lector revise el glosario de jerga antes de continuar con este artículo sobre reentrada atmosférica.
Cuando la entrada atmosférica es parte del aterrizaje o la recuperación de una nave espacial, particularmente en un cuerpo planetario que no sea la Tierra, la entrada es parte de una fase denominada entrada, descenso y aterrizaje o EDL. [7] Cuando la entrada atmosférica regresa al mismo cuerpo desde el que se había lanzado el vehículo, el evento se denomina reentrada (casi siempre se refiere a la entrada a la Tierra).
El objetivo de diseño fundamental en la entrada atmosférica de una nave espacial es disipar la energía de una nave espacial que viaja a velocidad hipersónica cuando ingresa a una atmósfera, de modo que el equipo, la carga y los pasajeros se ralentizan y aterrizan cerca de un destino específico en la superficie en velocidad cero mientras se mantienen las tensiones en la nave espacial y los pasajeros dentro de límites aceptables. [8] Esto puede lograrse mediante medios de propulsión o aerodinámicos (características del vehículo o paracaídas ), o mediante alguna combinación.
Formas de vehículos de entrada
Hay varias formas básicas que se utilizan en el diseño de vehículos de entrada:
Esfera o sección esférica
La forma axisimétrica más simple es la esfera o sección esférica. [9] Esto puede ser una esfera completa o un cuerpo anterior de sección esférica con un cuerpo posterior cónico convergente. La aerodinámica de una esfera o sección esférica es fácil de modelar analíticamente utilizando la teoría de impacto newtoniana. Asimismo, el flujo de calor de la sección esférica se puede modelar con precisión con la ecuación de Fay-Riddell. [10] La estabilidad estática de una sección esférica está asegurada si el centro de masa del vehículo está aguas arriba del centro de curvatura (la estabilidad dinámica es más problemática). Las esferas puras no tienen sustentación. Sin embargo, al volar en un ángulo de ataque , una sección esférica tiene una sustentación aerodinámica modesta, lo que proporciona cierta capacidad de rango transversal y ensancha su pasillo de entrada. A finales de la década de 1950 y principios de la de 1960, las computadoras de alta velocidad aún no estaban disponibles y la dinámica de fluidos computacional aún era embrionaria. Debido a que la sección esférica era susceptible de análisis de forma cerrada, esa geometría se convirtió en la predeterminada para el diseño conservador. En consecuencia, las cápsulas tripuladas de esa época se basaron en la sección esférica.
Los vehículos de entrada esférica pura se utilizaron en las primeras cápsulas soviéticas Vostok y Voskhod y en los vehículos de descenso soviéticos de Marte y Venera . El módulo de comando del Apolo utilizaba un escudo térmico de sección esférica en el cuerpo delantero con un cuerpo posterior cónico convergente. Voló una entrada de elevación con un ángulo de ataque hipersónico de -27 ° (0 ° es el extremo romo primero) para producir un L / D promedio (relación de elevación a arrastre) de 0.368. [11] El levantamiento resultante logró una medida de control de rango transversal al desviar el centro de masa del vehículo de su eje de simetría, permitiendo que la fuerza de levantamiento se dirija hacia la izquierda o hacia la derecha haciendo rodar la cápsula sobre su eje longitudinal . Otros ejemplos de la geometría de la sección esférica en cápsulas tripuladas son Soyuz / Zond , Gemini y Mercury . Incluso estas pequeñas cantidades de sustentación permiten trayectorias que tienen efectos muy significativos en la fuerza g máxima , reduciéndola de 8 a 9 g para una trayectoria puramente balística (ralentizada solo por arrastre) a 4 a 5 g, además de reducir en gran medida el pico. calor de reentrada. [12]
Esfera-cono
El cono esférico es una sección esférica con un cono truncado o desafilado adjunto. La estabilidad dinámica del cono esférico suele ser mejor que la de una sección esférica. El vehículo entra en la esfera primero. Con un semiángulo suficientemente pequeño y un centro de masa correctamente colocado, un cono esférico puede proporcionar estabilidad aerodinámica desde la entrada kepleriana hasta el impacto en la superficie. (La mitad del ángulo es el ángulo entre el eje de simetría rotacional del cono y su superficie exterior y, por lo tanto, la mitad del ángulo formado por los bordes de la superficie del cono).
El aeroshell de cono esférico estadounidense original fue el Mk-2 RV (vehículo de reentrada), que fue desarrollado en 1955 por General Electric Corp. El diseño del Mk-2 se derivó de la teoría del cuerpo contundente y utilizó un sistema de protección térmica refrigerado por radiación ( TPS) basado en un escudo térmico metálico (los diferentes tipos de TPS se describen más adelante en este artículo). El Mk-2 tenía defectos significativos como sistema de lanzamiento de armas, es decir, merodeaba demasiado tiempo en la atmósfera superior debido a su coeficiente balístico más bajo y también arrastraba una corriente de metal vaporizado haciéndolo muy visible para el radar . Estos defectos hicieron que el Mk-2 fuera demasiado susceptible a los sistemas de misiles antibalísticos (ABM). En consecuencia, General Electric desarrolló un RV de cono esférico alternativo al Mk-2. [ cita requerida ]
Este nuevo RV fue el Mk-6 que utilizó un TPS ablativo no metálico, un fenólico de nailon. Este nuevo TPS fue tan efectivo como un escudo térmico de reentrada que redujo significativamente la brusquedad. [ cita requerida ] Sin embargo, el Mk-6 era un enorme RV con una masa de entrada de 3360 kg, una longitud de 3,1 my un medio ángulo de 12,5 °. Los avances posteriores en el diseño de armas nucleares y TPS ablativo permitieron que los vehículos recreativos se volvieran significativamente más pequeños con una relación de contundencia aún más reducida en comparación con el Mk-6. Desde la década de 1960, el cono esférico se ha convertido en la geometría preferida para los RVs de misiles balísticos intercontinentales modernos con semiangulares típicos entre 10 ° y 11 °. [ cita requerida ]
Los vehículos recreativos (vehículos de recuperación) satelitales de reconocimiento también usaban una forma de cono esférico y fueron el primer ejemplo estadounidense de un vehículo de entrada sin municiones ( Discoverer-I , lanzado el 28 de febrero de 1959). El cono esférico se utilizó más tarde para misiones de exploración espacial a otros cuerpos celestes o para regresar desde el espacio abierto; por ejemplo, sonda Stardust . A diferencia de los vehículos recreativos militares, la ventaja de la masa TPS más baja del cuerpo contundente se mantuvo con los vehículos de entrada de exploración espacial como la sonda Galileo con un medio ángulo de 45 ° o el aeroshell Viking con un medio ángulo de 70 °. Los vehículos de exploración espacial con entrada de cono de esfera han aterrizado en la superficie o han entrado en las atmósferas de Marte , Venus , Júpiter y Titán .
Bicónico
El bicónico es un cono esférico con un tronco de árbol adicional adjunto. El biconic ofrece una relación L / D significativamente mejorada. Un bicónico diseñado para la aerocaptura de Marte generalmente tiene un L / D de aproximadamente 1.0 en comparación con un L / D de 0.368 para el Apollo-CM. La L / D más alta hace que una forma bicónica sea más adecuada para transportar personas a Marte debido a la desaceleración máxima más baja. Podría decirse que el bicónico más importante jamás volado fue el Vehículo de reentrada maniobrable avanzado (AMaRV). McDonnell Douglas Corp. fabricó cuatro AMaRV y representaron un salto significativo en la sofisticación de los vehículos recreativos. Los misiles balísticos intercontinentales Minuteman-1 lanzaron tres AMaRV el 20 de diciembre de 1979, el 8 de octubre de 1980 y el 4 de octubre de 1981. El AMaRV tenía una masa de entrada de aproximadamente 470 kg, un radio de morro de 2,34 cm, un semiangulo delantero y tronco de 10,4 °, un radio inter-frustum de 14,6 cm, medio ángulo de popa-frustum de 6 ° y una longitud axial de 2,079 metros. Ningún diagrama o imagen precisa de AMaRV ha aparecido nunca en la literatura abierta. Sin embargo, se ha publicado un boceto esquemático de un vehículo similar a AMaRV junto con diagramas de trayectoria que muestran curvas cerradas. [13]
La actitud de AMaRV se controló a través de una aleta de cuerpo dividida (también llamada aleta de barlovento dividida ) junto con dos aletas de guiñada montadas en los lados del vehículo. Se utilizó accionamiento hidráulico para controlar las aletas. AMaRV fue guiado por un sistema de navegación totalmente autónomo diseñado para evadir la interceptación de misiles antibalísticos (ABM). El McDonnell Douglas DC-X (también bicónico) era esencialmente una versión ampliada de AMaRV. AMaRV y el DC-X también sirvieron como base para una propuesta infructuosa de lo que finalmente se convirtió en el Lockheed Martin X-33 .
Formas no simétricas
Se han utilizado formas no simétricas para los vehículos de entrada tripulados. Un ejemplo es el vehículo en órbita alada que utiliza un ala delta para maniobrar durante el descenso de forma muy similar a un planeador convencional. Este enfoque ha sido utilizado por el transbordador espacial estadounidense y el Buran soviético . El cuerpo de elevación es otra geometría de vehículo de entrada y se usó con el vehículo X-23 PRIME (recuperación de precisión que incluye entrada de maniobra). [ cita requerida ]
Calefacción de reentrada
Los objetos que ingresan a la atmósfera desde el espacio a altas velocidades en relación con la atmósfera causarán niveles muy altos de calentamiento . El calentamiento de reentrada proviene principalmente de dos fuentes: [14]
- calefacción por convección , de dos tipos:
- El flujo de gas caliente pasa por la superficie del cuerpo y
- Reacciones de recombinación química catalítica entre la superficie del objeto y los gases atmosféricos.
- Calentamiento radiativo , desde la capa de choque energético que se forma al frente y a los lados del objeto.
A medida que aumenta la velocidad, aumentan tanto el calentamiento convectivo como el radiativo. A velocidades muy altas, el calentamiento radiativo llegará a dominar rápidamente los flujos de calor convectivo, ya que el calentamiento convectivo es proporcional a la velocidad al cubo, mientras que el calentamiento radiativo es proporcional a la octava potencia de velocidad. El calentamiento radiativo, que depende en gran medida de la longitud de onda , predomina muy temprano en la entrada atmosférica, mientras que la convección predomina en las últimas fases. [14]
Física del gas de la capa de choque
A temperaturas de reentrada típicas, el aire en la capa de choque está ionizado y disociado . [ cita requerida ] [15] Esta disociación química necesita varios modelos físicos para describir las propiedades químicas y térmicas de la capa de choque. Hay cuatro modelos físicos básicos de un gas que son importantes para los ingenieros aeronáuticos que diseñan escudos térmicos:
Modelo de gas perfecto
A casi todos los ingenieros aeronáuticos se les enseña el modelo de gas perfecto (ideal) durante su educación universitaria. La mayoría de las ecuaciones de gases perfectos importantes junto con sus tablas y gráficos correspondientes se muestran en el Informe NACA 1135. [16] Los extractos del Informe NACA 1135 a menudo aparecen en los apéndices de los libros de texto de termodinámica y son familiares para la mayoría de los ingenieros aeronáuticos que diseñan aviones supersónicos.
La teoría del gas perfecto es elegante y extremadamente útil para diseñar aviones, pero asume que el gas es químicamente inerte. Desde el punto de vista del diseño de aeronaves, se puede suponer que el aire es inerte a temperaturas inferiores a 550 K a una presión atmosférica. La teoría del gas perfecto comienza a descomponerse a 550 K y no se puede utilizar a temperaturas superiores a 2000 K. Para temperaturas superiores a 2000 K, un diseñador de protección térmica debe utilizar un modelo de gas real .
Modelo de gas real (de equilibrio)
El momento de cabeceo de un vehículo de entrada puede verse significativamente influenciado por los efectos del gas real. Tanto el módulo de comando Apollo como el transbordador espacial se diseñaron utilizando momentos de cabeceo incorrectos determinados a través de un modelado inexacto de gas real. El ángulo de ataque del ángulo de compensación del Apollo-CM fue más alto de lo estimado originalmente, lo que resultó en un corredor de entrada de retorno lunar más estrecho. El centro aerodinámico real del Columbia estaba aguas arriba del valor calculado debido a los efectos del gas real. En Columbia ' vuelo inaugural s ( STS-1 ), los astronautas John W. Young y Robert Crippen tenían algunos momentos de ansiedad durante la reentrada, cuando existía la preocupación por la pérdida de control del vehículo. [17]
Un modelo de equilibrio de gas real supone que un gas es químicamente reactivo, pero también supone que todas las reacciones químicas han tenido tiempo de completarse y que todos los componentes del gas tienen la misma temperatura (esto se llama equilibrio termodinámico ). Cuando el aire es procesado por una onda de choque, se sobrecalienta por compresión y se disocia químicamente a través de muchas reacciones diferentes. La fricción directa sobre el objeto de reentrada no es la principal causa del calentamiento de la capa de choque. Se debe principalmente al calentamiento isentrópico de las moléculas de aire dentro de la onda de compresión. Los aumentos de entropía basados en la fricción de las moléculas dentro de la onda también explican algo de calentamiento. [ investigación original? ] La distancia desde la onda de choque al punto de estancamiento en el borde más importantes del vehículo de entrada se denomina onda de choque stand off . Una regla general aproximada para la distancia de separación de la onda de choque es 0,14 veces el radio de la punta. Se puede estimar el tiempo de viaje de una molécula de gas desde la onda de choque hasta el punto de estancamiento asumiendo una velocidad de flujo libre de 7.8 km / sy un radio de punta de 1 metro, es decir, el tiempo de viaje es de aproximadamente 18 microsegundos. Este es aproximadamente el tiempo necesario para que la disociación química iniciada por ondas de choque se acerque al equilibrio químico en una capa de choque para una entrada de 7,8 km / s en el aire durante el flujo de calor máximo. En consecuencia, a medida que el aire se acerca al punto de estancamiento del vehículo de entrada, el aire alcanza efectivamente el equilibrio químico, lo que permite utilizar un modelo de equilibrio. Para este caso, la mayor parte de la capa de choque entre la onda de choque y el borde de ataque de un vehículo de entrada está reaccionando químicamente y no en un estado de equilibrio. La ecuación de Fay-Riddell , [10] que es de extrema importancia para modelar el flujo de calor, debe su validez al punto de estancamiento que se encuentra en equilibrio químico. El tiempo necesario para que el gas de la capa de choque alcance el equilibrio depende en gran medida de la presión de la capa de choque. Por ejemplo, en el caso de la entrada de la sonda Galileo en la atmósfera de Júpiter, la capa de choque estaba mayormente en equilibrio durante el pico de flujo de calor debido a las muy altas presiones experimentadas (esto es contrario a la intuición dado que la velocidad de la corriente libre era de 39 km / s durante el pico de calor). flujo).
Determinar el estado termodinámico del punto de estancamiento es más difícil en un modelo de gas de equilibrio que en un modelo de gas perfecto. En un modelo de gas perfecto, se supone que la relación de calores específicos (también denominada exponente isentrópico , índice adiabático , gamma o kappa ) es constante junto con la constante del gas . Para un gas real, la proporción de calores específicos puede oscilar enormemente en función de la temperatura. Bajo un modelo de gas perfecto, existe un elegante conjunto de ecuaciones para determinar el estado termodinámico a lo largo de una línea de corriente de entropía constante llamada cadena isentrópica . Para un gas real, la cadena isentrópica no se puede utilizar y, en su lugar, se utilizaría un diagrama de Mollier para el cálculo manual. Sin embargo, la solución gráfica con un diagrama de Mollier ahora se considera obsoleta con los diseñadores modernos de escudos térmicos que utilizan programas de computadora basados en una tabla de búsqueda digital (otra forma de diagrama de Mollier) o un programa de termodinámica basado en la química. La composición química de un gas en equilibrio con presión y temperatura fijas se puede determinar mediante el método de energía libre de Gibbs . La energía libre de Gibbs es simplemente la entalpía total del gas menos su entropía total multiplicada por la temperatura. Un programa de equilibrio químico normalmente no requiere fórmulas químicas o ecuaciones de velocidad de reacción. El programa funciona preservando las abundancias elementales originales especificadas para el gas y variando las diferentes combinaciones moleculares de los elementos mediante iteración numérica hasta que se calcula la energía libre de Gibbs más baja posible (un método de Newton-Raphson es el esquema numérico habitual). La base de datos para un programa de energía libre de Gibbs proviene de datos espectroscópicos utilizados para definir funciones de partición . Entre los mejores códigos de equilibrio que existen está el programa Equilibrio químico con aplicaciones (CEA) que fue escrito por Bonnie J. McBride y Sanford Gordon en NASA Lewis (ahora rebautizado como "Centro de Investigación Glenn de la NASA"). Otros nombres de CEA son el "Código Gordon y McBride" y el "Código Lewis". CEA es bastante preciso hasta 10,000 K para gases atmosféricos planetarios, pero inutilizable más allá de 20,000 K ( no se modela la doble ionización ). CEA se puede descargar de Internet junto con la documentación completa y se compilará en Linux bajo el compilador G77 Fortran .
Modelo de gas real (sin equilibrio)
Un modelo de gas real de no equilibrio es el modelo más preciso de la física del gas de una capa de choque, pero es más difícil de resolver que un modelo de equilibrio. El modelo de no equilibrio más simple es el modelo Lighthill-Freeman desarrollado en 1958. [18] [19] El modelo Lighthill-Freeman inicialmente asume un gas compuesto por una sola especie diatómica susceptible a una sola fórmula química y su inversa; por ejemplo, N 2 ? N + N y N + N? N 2 (disociación y recombinación). Debido a su simplicidad, el modelo Lighthill-Freeman es una herramienta pedagógica útil, pero desafortunadamente es demasiado simple para modelar el aire en desequilibrio. Normalmente se supone que el aire tiene una composición de fracción molar de 0,7812 nitrógeno molecular, 0,2095 oxígeno molecular y 0,0093 argón. El modelo de gas real más simple para el aire es el modelo de cinco especies , que se basa en N 2 , O 2 , NO, N y O. El modelo de cinco especies asume que no hay ionización e ignora trazas de especies como el dióxido de carbono.
Cuando se ejecuta un programa de equilibrio de energía libre de Gibbs, [ aclaración necesaria ] el proceso iterativo desde la composición molecular originalmente especificada hasta la composición de equilibrio calculada final es esencialmente aleatorio y no preciso en el tiempo. Con un programa de no equilibrio, el proceso de cálculo es preciso en el tiempo y sigue una ruta de solución dictada por fórmulas químicas y de velocidad de reacción. El modelo de cinco especies tiene 17 fórmulas químicas (34 cuando se cuentan fórmulas inversas). El modelo de Lighthill-Freeman se basa en una única ecuación diferencial ordinaria y una ecuación algebraica. El modelo de cinco especies se basa en 5 ecuaciones diferenciales ordinarias y 17 ecuaciones algebraicas. [ cita requerida ] Debido a que las 5 ecuaciones diferenciales ordinarias están estrechamente acopladas, el sistema es numéricamente "rígido" y difícil de resolver. El modelo de cinco especies solo se puede utilizar para la entrada desde la órbita terrestre baja, donde la velocidad de entrada es de aproximadamente 7,8 km / s (28.000 km / h; 17.000 mph). Para una entrada de retorno lunar de 11 km / s, [20] la capa de choque contiene una cantidad significativa de nitrógeno y oxígeno ionizados. El modelo de cinco especies ya no es exacto y en su lugar se debe utilizar un modelo de doce especies. Las velocidades de la interfaz de entrada atmosférica [se necesita aclaración ] en una trayectoria Marte-Tierra son del orden de 12 km / s (43.000 km / h; 27.000 mph). [21] Modelar la entrada atmosférica de Marte a alta velocidad, que involucra una atmósfera de dióxido de carbono, nitrógeno y argón, es aún más complejo y requiere un modelo de 19 especies. [ cita requerida ]
An important aspect of modelling non-equilibrium real gas effects is radiative heat flux. If a vehicle is entering an atmosphere at very high speed (hyperbolic trajectory, lunar return) and has a large nose radius then radiative heat flux can dominate TPS heating. Radiative heat flux during entry into an air or carbon dioxide atmosphere typically comes from asymmetric diatomic molecules; e.g., cyanogen (CN), carbon monoxide, nitric oxide (NO), single ionized molecular nitrogen etc. These molecules are formed by the shock wave dissociating ambient atmospheric gas followed by recombination within the shock layer into new molecular species. The newly formed diatomic molecules initially have a very high vibrational temperature that efficiently transforms the vibrational energy into radiant energy; i.e., radiative heat flux. The whole process takes place in less than a millisecond which makes modelling a challenge. The experimental measurement of radiative heat flux (typically done with shock tubes) along with theoretical calculation through the unsteady Schrödinger equation are among the more esoteric aspects of aerospace engineering. Most of the aerospace research work related to understanding radiative heat flux was done in the 1960s, but largely discontinued after conclusion of the Apollo Program. Radiative heat flux in air was just sufficiently understood to ensure Apollo's success. However, radiative heat flux in carbon dioxide (Mars entry) is still barely understood and will require major research.[citation needed]
Frozen gas model
The frozen gas model describes a special case of a gas that is not in equilibrium. The name "frozen gas" can be misleading. A frozen gas is not "frozen" like ice is frozen water. Rather a frozen gas is "frozen" in time (all chemical reactions are assumed to have stopped). Chemical reactions are normally driven by collisions between molecules. If gas pressure is slowly reduced such that chemical reactions can continue then the gas can remain in equilibrium. However, it is possible for gas pressure to be so suddenly reduced that almost all chemical reactions stop. For that situation the gas is considered frozen.[citation needed]
The distinction between equilibrium and frozen is important because it is possible for a gas such as air to have significantly different properties (speed-of-sound, viscosity etc.) for the same thermodynamic state; e.g., pressure and temperature. Frozen gas can be a significant issue in the wake behind an entry vehicle. During reentry, free stream air is compressed to high temperature and pressure by the entry vehicle's shock wave. Non-equilibrium air in the shock layer is then transported past the entry vehicle's leading side into a region of rapidly expanding flow that causes freezing. The frozen air can then be entrained into a trailing vortex behind the entry vehicle. Correctly modelling the flow in the wake of an entry vehicle is very difficult. Thermal protection shield (TPS) heating in the vehicle's afterbody is usually not very high, but the geometry and unsteadiness of the vehicle's wake can significantly influence aerodynamics (pitching moment) and particularly dynamic stability.[citation needed]
Sistemas de protección térmica
A thermal protection system, or TPS, is the barrier that protects a spacecraft during the searing heat of atmospheric reentry. A secondary goal may be to protect the spacecraft from the heat and cold of space while in orbit. Multiple approaches for the thermal protection of spacecraft are in use, among them ablative heat shields, passive cooling, and active cooling of spacecraft surfaces.
Ablative
The ablative heat shield functions by lifting the hot shock layer gas away from the heat shield's outer wall (creating a cooler boundary layer). The boundary layer comes from blowing of gaseous reaction products from the heat shield material and provides protection against all forms of heat flux. The overall process of reducing the heat flux experienced by the heat shield's outer wall by way of a boundary layer is called blockage. Ablation occurs at two levels in an ablative TPS: the outer surface of the TPS material chars, melts, and sublimes, while the bulk of the TPS material undergoes pyrolysis and expels product gases. The gas produced by pyrolysis is what drives blowing and causes blockage of convective and catalytic heat flux. Pyrolysis can be measured in real time using thermogravimetric analysis, so that the ablative performance can be evaluated.[22] Ablation can also provide blockage against radiative heat flux by introducing carbon into the shock layer thus making it optically opaque. Radiative heat flux blockage was the primary thermal protection mechanism of the Galileo Probe TPS material (carbon phenolic). Carbon phenolic was originally developed as a rocket nozzle throat material (used in the Space Shuttle Solid Rocket Booster) and for reentry-vehicle nose tips.
Early research on ablation technology in the USA was centered at NASA's Ames Research Center located at Moffett Field, California. Ames Research Center was ideal, since it had numerous wind tunnels capable of generating varying wind velocities. Initial experiments typically mounted a mock-up of the ablative material to be analyzed within a hypersonic wind tunnel.[23] Testing of ablative materials occurs at the Ames Arc Jet Complex. Many spacecraft thermal protection systems have been tested in this facility, including the Apollo, space shuttle, and Orion heat shield materials.[24]
The thermal conductivity of a particular TPS material is usually proportional to the material's density.[25] Carbon phenolic is a very effective ablative material, but also has high density which is undesirable. If the heat flux experienced by an entry vehicle is insufficient to cause pyrolysis then the TPS material's conductivity could allow heat flux conduction into the TPS bondline material thus leading to TPS failure. Consequently, for entry trajectories causing lower heat flux, carbon phenolic is sometimes inappropriate and lower-density TPS materials such as the following examples can be better design choices:
Super light-weight ablator
SLA in SLA-561V stands for super light-weight ablator. SLA-561V is a proprietary ablative made by Lockheed Martin that has been used as the primary TPS material on all of the 70° sphere-cone entry vehicles sent by NASA to Mars other than the Mars Science Laboratory (MSL). SLA-561V begins significant ablation at a heat flux of approximately 110 W/cm2, but will fail for heat fluxes greater than 300 W/cm2. The MSL aeroshell TPS is currently designed to withstand a peak heat flux of 234 W/cm2. The peak heat flux experienced by the Viking 1 aeroshell which landed on Mars was 21 W/cm2. For Viking 1, the TPS acted as a charred thermal insulator and never experienced significant ablation. Viking 1 was the first Mars lander and based upon a very conservative design. The Viking aeroshell had a base diameter of 3.54 meters (the largest used on Mars until Mars Science Laboratory). SLA-561V is applied by packing the ablative material into a honeycomb core that is pre-bonded to the aeroshell's structure thus enabling construction of a large heat shield.[26]
Phenolic-impregnated carbon ablator
Phenolic-impregnated carbon ablator (PICA), a carbon fiber preform impregnated in phenolic resin,[27] is a modern TPS material and has the advantages of low density (much lighter than carbon phenolic) coupled with efficient ablative ability at high heat flux. It is a good choice for ablative applications such as high-peak-heating conditions found on sample-return missions or lunar-return missions. PICA's thermal conductivity is lower than other high-heat-flux-ablative materials, such as conventional carbon phenolics.[citation needed]
PICA was patented by NASA Ames Research Center in the 1990s and was the primary TPS material for the Stardust aeroshell.[28] The Stardust sample-return capsule was the fastest man-made object ever to reenter Earth's atmosphere (12.4 km/s (28,000 mph) at 135 km altitude). This was faster than the Apollo mission capsules and 70% faster than the Shuttle.[29] PICA was critical for the viability of the Stardust mission, which returned to Earth in 2006. Stardust's heat shield (0.81 m base diameter) was made of one monolithic piece sized to withstand a nominal peak heating rate of 1.2 kW/cm2. A PICA heat shield was also used for the Mars Science Laboratory entry into the Martian atmosphere.[30]
PICA-X
An improved and easier to produce version called PICA-X was developed by SpaceX in 2006–2010[30] for the Dragon space capsule.[31] The first reentry test of a PICA-X heat shield was on the Dragon C1 mission on 8 December 2010.[32] The PICA-X heat shield was designed, developed and fully qualified by a small team of a dozen engineers and technicians in less than four years.[30] PICA-X is ten times less expensive to manufacture than the NASA PICA heat shield material.[33]
PICA-3
A second enhanced version of PICA—called PICA-3—was developed by SpaceX during the mid-2010s. It was first flight tested on the Crew Dragon spacecraft in 2019 during the flight demonstration mission, in April 2019, and put into regular service on that spacecraft in 2020.[34]
SIRCA
Silicone-impregnated reusable ceramic ablator (SIRCA) was also developed at NASA Ames Research Center and was used on the Backshell Interface Plate (BIP) of the Mars Pathfinder and Mars Exploration Rover (MER) aeroshells. The BIP was at the attachment points between the aeroshell's backshell (also called the afterbody or aft cover) and the cruise ring (also called the cruise stage). SIRCA was also the primary TPS material for the unsuccessful Deep Space 2 (DS/2) Mars impactor probes with their 0.35-meter-base-diameter (1.1 ft) aeroshells. SIRCA is a monolithic, insulating material that can provide thermal protection through ablation. It is the only TPS material that can be machined to custom shapes and then applied directly to the spacecraft. There is no post-processing, heat treating, or additional coatings required (unlike Space Shuttle tiles). Since SIRCA can be machined to precise shapes, it can be applied as tiles, leading edge sections, full nose caps, or in any number of custom shapes or sizes. As of 1996[update], SIRCA had been demonstrated in backshell interface applications, but not yet as a forebody TPS material.[35]
AVCOAT
AVCOAT is a NASA-specified ablative heat shield, a glass-filled epoxy–novolac system.[36]
NASA originally used it for the Apollo capsule in the 1960s, and then utilized the material for its next-generation beyond low-Earth-orbit Orion spacecraft, slated to fly in the early 2020s.[37] The Avcoat to be used on Orion has been reformulated to meet environmental legislation that has been passed since the end of Apollo.[38][39]
Thermal soak
Thermal soak is a part of almost all TPS schemes. For example, an ablative heat shield loses most of its thermal protection effectiveness when the outer wall temperature drops below the minimum necessary for pyrolysis. From that time to the end of the heat pulse, heat from the shock layer convects into the heat shield's outer wall and would eventually conduct to the payload.[citation needed] This outcome is prevented by ejecting the heat shield (with its heat soak) prior to the heat conducting to the inner wall.
Typical Space Shuttle TPS tiles (LI-900) have remarkable thermal protection properties. An LI-900 tile exposed to a temperature of 1,000 K on one side will remain merely warm to the touch on the other side. However, they are relatively brittle and break easily, and cannot survive in-flight rain.
Passively cooled
In some early ballistic missile RVs (e.g., the Mk-2 and the sub-orbital Mercury spacecraft), radiatively cooled TPS were used to initially absorb heat flux during the heat pulse, and, then, after the heat pulse, radiate and convect the stored heat back into the atmosphere. However, the earlier version of this technique required a considerable quantity of metal TPS (e.g., titanium, beryllium, copper, etc.). Modern designers prefer to avoid this added mass by using ablative and thermal-soak TPS instead.
Thermal protection systems relying on emissivity use high emissivity coatings (HECs) to facilitate radiative cooling, while an underlying porous ceramic layer serves to protect the structure from high surface temperatures. High thermally stable emissivity values coupled with low thermal conductivity are key to the functionality of such systems.[40]
Radiatively cooled TPS can be found on modern entry vehicles, but reinforced carbon–carbon (RCC) (also called carbon–carbon) is normally used instead of metal. RCC was the TPS material on the Space Shuttle's nose cone and wing leading edges, and was also proposed as the leading-edge material for the X-33. Carbon is the most refractory material known, with a one-atmosphere sublimation temperature of 3,825 °C (6,917 °F) for graphite. This high temperature made carbon an obvious choice as a radiatively cooled TPS material. Disadvantages of RCC are that it is currently expensive to manufacture, is heavy, and lacks robust impact resistance.[41]
Some high-velocity aircraft, such as the SR-71 Blackbird and Concorde, deal with heating similar to that experienced by spacecraft, but at much lower intensity, and for hours at a time. Studies of the SR-71's titanium skin revealed that the metal structure was restored to its original strength through annealing due to aerodynamic heating. In the case of the Concorde, the aluminium nose was permitted to reach a maximum operating temperature of 127 °C (261 °F) (approximately 180 °C (324 °F) warmer than the normally sub-zero, ambient air); the metallurgical implications (loss of temper) that would be associated with a higher peak temperature were the most significant factors determining the top speed of the aircraft.
A radiatively cooled TPS for an entry vehicle is often called a hot-metal TPS. Early TPS designs for the Space Shuttle called for a hot-metal TPS based upon a nickel superalloy (dubbed René 41) and titanium shingles.[42] This Shuttle TPS concept was rejected, because it was believed a silica tile-based TPS would involve lower development and manufacturing costs.[citation needed] A nickel superalloy-shingle TPS was again proposed for the unsuccessful X-33 single-stage-to-orbit (SSTO) prototype.[43]
Recently, newer radiatively cooled TPS materials have been developed that could be superior to RCC. Known as Ultra-High Temperature Ceramics, they were developed for the prototype vehicle Slender Hypervelocity Aerothermodynamic Research Probe (SHARP). These TPS materials are based on zirconium diboride and hafnium diboride. SHARP TPS have suggested performance improvements allowing for sustained Mach 7 flight at sea level, Mach 11 flight at 100,000-foot (30,000 m) altitudes, and significant improvements for vehicles designed for continuous hypersonic flight. SHARP TPS materials enable sharp leading edges and nose cones to greatly reduce drag for airbreathing combined-cycle-propelled spaceplanes and lifting bodies. SHARP materials have exhibited effective TPS characteristics from zero to more than 2,000 °C (3,630 °F), with melting points over 3,500 °C (6,330 °F). They are structurally stronger than RCC, and, thus, do not require structural reinforcement with materials such as Inconel. SHARP materials are extremely efficient at reradiating absorbed heat, thus eliminating the need for additional TPS behind and between the SHARP materials and conventional vehicle structure. NASA initially funded (and discontinued) a multi-phase R&D program through the University of Montana in 2001 to test SHARP materials on test vehicles.[44][45]
Actively cooled
Various advanced reusable spacecraft and hypersonic aircraft designs have been proposed to employ heat shields made from temperature-resistant metal alloys that incorporate a refrigerant or cryogenic fuel circulating through them, and one such spacecraft design is currently under development.
Such a TPS concept was proposed[when?] for the X-30 National Aerospace Plane (NASP).[citation needed] The NASP was supposed to have been a scramjet powered hypersonic aircraft, but failed in development.
SpaceX is currently developing an actively cooled heat shield for its Starship spacecraft where a part of the thermal protection system will be a transpirationally cooled outer-skin design for the reentering spaceship.[46][47]
In the early 1960s various TPS systems were proposed to use water or other cooling liquid sprayed into the shock layer, or passed through channels in the heat shield. Advantages included the possibility of more all-metal designs which would be cheaper to develop, be more rugged, and eliminate the need for classified technology. The disadvantages are increased weight and complexity, and lower reliability. The concept has never been flown, but a similar technology (the plug nozzle[48]) did undergo extensive ground testing.
Reentrada emplumada
In 2004, aircraft designer Burt Rutan demonstrated the feasibility of a shape-changing airfoil for reentry with the sub-orbital SpaceShipOne. The wings on this craft rotate upward into the feathered configuration that provides a shuttlecock effect. Thus SpaceShipOne achieves much more aerodynamic drag on reentry while not experiencing significant thermal loads.
The configuration increases drag, as the craft is now less streamlined and results in more atmospheric gas particles hitting the spacecraft at higher altitudes than otherwise. The aircraft thus slows down more in higher atmospheric layers which is the key to efficient reentry. Secondly, the aircraft will automatically orient itself in this state to a high drag attitude.[49]
However, the velocity attained by SpaceShipOne prior to reentry is much lower than that of an orbital spacecraft, and engineers, including Rutan, recognize that a feathered reentry technique is not suitable for return from orbit.
On 4 May 2011, the first test on the SpaceShipTwo of the feathering mechanism was made during a glideflight after release from the White Knight Two. Premature deployment of the feathering system was responsible for the 2014 VSS Enterprise crash, in which the aircraft disintegrated, killing the co-pilot.
The feathered reentry was first described by Dean Chapman of NACA in 1958.[50] In the section of his report on Composite Entry, Chapman described a solution to the problem using a high-drag device:
It may be desirable to combine lifting and nonlifting entry in order to achieve some advantages... For landing maneuverability it obviously is advantageous to employ a lifting vehicle. The total heat absorbed by a lifting vehicle, however, is much higher than for a nonlifting vehicle... Nonlifting vehicles can more easily be constructed... by employing, for example, a large, light drag device... The larger the device, the smaller is the heating rate.
Nonlifting vehicles with shuttlecock stability are advantageous also from the viewpoint of minimum control requirements during entry.
... an evident composite type of entry, which combines some of the desirable features of lifting and nonlifting trajectories, would be to enter first without lift but with a... drag device; then, when the velocity is reduced to a certain value... the device is jettisoned or retracted, leaving a lifting vehicle... for the remainder of the descent.
The North American X-15 used a similar mechanism.[citation needed]
Reentrada con escudo térmico inflable
Deceleration for atmospheric reentry, especially for higher-speed Mars-return missions, benefits from maximizing "the drag area of the entry system. The larger the diameter of the aeroshell, the bigger the payload can be."[51] An inflatable aeroshell provides one alternative for enlarging the drag area with a low-mass design.
Non-US
Such an inflatable shield/aerobrake was designed for the penetrators of Mars 96 mission. Since the mission failed due to the launcher malfunction, the NPO Lavochkin and DASA/ESA have designed a mission for Earth orbit. The Inflatable Reentry and Descent Technology (IRDT) demonstrator was launched on Soyuz-Fregat on 8 February 2000. The inflatable shield was designed as a cone with two stages of inflation. Although the second stage of the shield failed to inflate, the demonstrator survived the orbital reentry and was recovered.[52][53] The subsequent missions flown on the Volna rocket failed due to launcher failure.[54]
NASA IRVE
NASA launched an inflatable heat shield experimental spacecraft on 17 August 2009 with the successful first test flight of the Inflatable Re-entry Vehicle Experiment (IRVE). The heat shield had been vacuum-packed into a 15-inch-diameter (38 cm) payload shroud and launched on a Black Brant 9 sounding rocket from NASA's Wallops Flight Facility on Wallops Island, Virginia. "Nitrogen inflated the 10-foot-diameter (3.0 m) heat shield, made of several layers of silicone-coated [Kevlar] fabric, to a mushroom shape in space several minutes after liftoff."[51] The rocket apogee was at an altitude of 131 miles (211 km) where it began its descent to supersonic speed. Less than a minute later the shield was released from its cover to inflate at an altitude of 124 miles (200 km). The inflation of the shield took less than 90 seconds.[51]
NASA HIAD
Following the success of the initial IRVE experiments, NASA developed the concept into the more ambitious Hypersonic Inflatable Aerodynamic Decelerator (HIAD). The current design is shaped like a shallow cone, with the structure built up as a stack of circular inflated tubes of gradually increasing major diameter. The forward (convex) face of the cone is covered with a flexible thermal protection system robust enough to withstand the stresses of atmospheric entry (or reentry).[55][56]
In 2012, a HIAD was tested as Inflatable Reentry Vehicle Experiment 3 (IRVE-3) using a sub-orbital sounding rocket, and worked.[57]:8
In 2020 there were plans to launch in 2022 a 6 m inflatable as Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID).[58]
See also Low-Density Supersonic Decelerator, a NASA project with tests in 2014 & 2015.
Consideraciones de diseño de vehículos de entrada
There are four critical parameters[according to whom?] considered when designing a vehicle for atmospheric entry:[citation needed]
- Peak heat flux
- Heat load
- Peak deceleration
- Peak dynamic pressure
Peak heat flux and dynamic pressure selects the TPS material. Heat load selects the thickness of the TPS material stack. Peak deceleration is of major importance for manned missions. The upper limit for manned return to Earth from low Earth orbit (LEO) or lunar return is 10g.[59] For Martian atmospheric entry after long exposure to zero gravity, the upper limit is 4g.[59] Peak dynamic pressure can also influence the selection of the outermost TPS material if spallation is an issue.
Starting from the principle of conservative design, the engineer typically considers two worst-case trajectories, the undershoot and overshoot trajectories. The overshoot trajectory is typically defined as the shallowest-allowable entry velocity angle prior to atmospheric skip-off. The overshoot trajectory has the highest heat load and sets the TPS thickness. The undershoot trajectory is defined by the steepest allowable trajectory. For manned missions the steepest entry angle is limited by the peak deceleration. The undershoot trajectory also has the highest peak heat flux and dynamic pressure. Consequently, the undershoot trajectory is the basis for selecting the TPS material. There is no "one size fits all" TPS material. A TPS material that is ideal for high heat flux may be too conductive (too dense) for a long duration heat load. A low-density TPS material might lack the tensile strength to resist spallation if the dynamic pressure is too high. A TPS material can perform well for a specific peak heat flux, but fail catastrophically for the same peak heat flux if the wall pressure is significantly increased (this happened with NASA's R-4 test spacecraft).[59] Older TPS materials tend to be more labor-intensive and expensive to manufacture compared to modern materials. However, modern TPS materials often lack the flight history of the older materials (an important consideration for a risk-averse designer).
Based upon Allen and Eggers discovery, maximum aeroshell bluntness (maximum drag) yields minimum TPS mass. Maximum bluntness (minimum ballistic coefficient) also yields a minimal terminal velocity at maximum altitude (very important for Mars EDL, but detrimental for military RVs). However, there is an upper limit to bluntness imposed by aerodynamic stability considerations based upon shock wave detachment. A shock wave will remain attached to the tip of a sharp cone if the cone's half-angle is below a critical value. This critical half-angle can be estimated using perfect gas theory (this specific aerodynamic instability occurs below hypersonic speeds). For a nitrogen atmosphere (Earth or Titan), the maximum allowed half-angle is approximately 60°. For a carbon dioxide atmosphere (Mars or Venus), the maximum-allowed half-angle is approximately 70°. After shock wave detachment, an entry vehicle must carry significantly more shocklayer gas around the leading edge stagnation point (the subsonic cap). Consequently, the aerodynamic center moves upstream thus causing aerodynamic instability. It is incorrect to reapply an aeroshell design intended for Titan entry (Huygens probe in a nitrogen atmosphere) for Mars entry (Beagle 2 in a carbon dioxide atmosphere).[citation needed][original research?] Prior to being abandoned, the Soviet Mars lander program achieved one successful landing (Mars 3), on the second of three entry attempts (the others were Mars 2 and Mars 6). The Soviet Mars landers were based upon a 60° half-angle aeroshell design.
A 45° half-angle sphere-cone is typically used for atmospheric probes (surface landing not intended) even though TPS mass is not minimized. The rationale for a 45° half-angle is to have either aerodynamic stability from entry-to-impact (the heat shield is not jettisoned) or a short-and-sharp heat pulse followed by prompt heat shield jettison. A 45° sphere-cone design was used with the DS/2 Mars impactor and Pioneer Venus probes.
Accidentes notables de entrada a la atmósfera
Not all atmospheric reentries have been successful and some have resulted in significant disasters.
- Voskhod 2 – The service module failed to detach for some time, but the crew survived.
- Soyuz 1 – The attitude control system failed while still in orbit and later parachutes got entangled during the emergency landing sequence (entry, descent, and landing (EDL) failure). Lone cosmonaut Vladimir Mikhailovich Komarov died.
- Soyuz 5 – The service module failed to detach, but the crew survived.
- Soyuz 11 – After tri-module separation, a valve was weakened by the blast and failed on reentry. The cabin depressurized killing all three crew members.
- Mars Polar Lander – Failed during EDL. The failure was believed to be the consequence of a software error. The precise cause is unknown for lack of real-time telemetry.
- Space Shuttle Columbia
- STS-1 – a combination of launch damage, protruding gap filler, and tile installation error resulted in serious damage to the orbiter, only some of which the crew was privy to. Had the crew known the true extent of the damage before attempting reentry, they would have flown the shuttle to a safe altitude and then bailed out. Nevertheless, reentry was successful, and the orbiter proceeded to a normal landing.
- STS-107 – The failure of an RCC panel on a wing leading edge caused by debris impact at launch led to breakup of the orbiter on reentry resulting in the deaths of all seven crew members.
- Genesis – The parachute failed to deploy due to a G-switch having been installed backwards (a similar error delayed parachute deployment for the Galileo Probe). Consequently, the Genesis entry vehicle crashed into the desert floor. The payload was damaged, but most scientific data were recoverable.
- Soyuz TMA-11 – The Soyuz propulsion module failed to separate properly; fallback ballistic reentry was executed that subjected the crew to accelerations of about 8 standard gravities (78 m/s2).[60] The crew survived.
Reentradas incontroladas y desprotegidas
Of satellites that reenter, approximately 10–40% of the mass of the object is likely to reach the surface of the Earth.[61] On average, about one catalogued object reenters per day.[62]
Due to the Earth's surface being primarily water, most objects that survive reentry land in one of the world's oceans. The estimated chances that a given person will get hit and injured during his/her lifetime is around 1 in a trillion.[63]
On January 24, 1978, the Soviet Kosmos 954 (3,800 kilograms [8,400 lb]) reentered and crashed near Great Slave Lake in the Northwest Territories of Canada. The satellite was nuclear-powered and left radioactive debris near its impact site.[64]
On July 11, 1979, the US Skylab space station (77,100 kilograms [170,000 lb]) reentered and spread debris across the Australian Outback.[65] The reentry was a major media event largely due to the Cosmos 954 incident, but not viewed as much as a potential disaster since it did not carry toxic nuclear or hydrazine fuel. NASA had originally hoped to use a Space Shuttle mission to either extend its life or enable a controlled reentry, but delays in the Shuttle program, plus unexpectedly high solar activity, made this impossible.[66][67]
On February 7, 1991, the Soviet Salyut 7 space station (19,820 kilograms [43,700 lb]), with the Kosmos 1686 module (20,000 kilograms [44,000 lb]) attached, reentered and scattered debris over the town of Capitán Bermúdez, Argentina.[68][69][70] The station had been boosted to a higher orbit in August 1986 in an attempt to keep it up until 1994, but in a scenario similar to Skylab, the planned Buran shuttle was cancelled and high solar activity caused it to come down sooner than expected.
On September 7, 2011, NASA announced the impending uncontrolled reentry of the Upper Atmosphere Research Satellite (6,540 kilograms [14,420 lb]) and noted that there was a small risk to the public.[71] The decommissioned satellite reentered the atmosphere on September 24, 2011, and some pieces are presumed to have crashed into the South Pacific Ocean over a debris field 500 miles (800 km) long.[72]
On April 1, 2018, the Chinese Tiangong-1 space station (8,510 kilograms [18,760 lb]) reentered over the Pacific Ocean, halfway between Australia and South America.[73] The China Manned Space Engineering Office had intended to control the reentry, but lost telemetry and control in March 2017.[74]
On May 11, 2020, the core stage of Chinese Long March 5B (COSPAR ID 2020-027C) weighing roughly 20,000 kilograms [44,000 lb]) made an uncontrolled reentry over the Atlantic Ocean, near West African coast.[75][76] Few pieces of rocket debris reportedly survived reentry and fell over at least two villages in Ivory Coast.[77][78]
It is expected that the Cruise Mass Balance Devices (CMBDs) from the Mars 2020 mission, which are ejected prior to the spacecraft entering the atmosphere, will survive re-entry and impact the surface on Thursday 18 February 2021.[79] The CMBDs are 77 kg tungsten blocks used to adjust the spacecraft's trajectory prior to entry. The Science Team of another NASA mission, InSight, announced in early 2021 that they would attempt to detect the seismic waves from this impact event.
Deorbit disposal
Salyut 1, the world's first space station, was deliberately de-orbited into the Pacific Ocean in 1971 following the Soyuz 11 accident. Its successor, Salyut 6, was de-orbited in a controlled manner as well.
On June 4, 2000 the Compton Gamma Ray Observatory was deliberately de-orbited after one of its gyroscopes failed. The debris that did not burn up fell harmlessly into the Pacific Ocean. The observatory was still operational, but the failure of another gyroscope would have made de-orbiting much more difficult and dangerous. With some controversy, NASA decided in the interest of public safety that a controlled crash was preferable to letting the craft come down at random.
In 2001, the Russian Mir space station was deliberately de-orbited, and broke apart in the fashion expected by the command center during atmospheric reentry. Mir entered the Earth's atmosphere on March 23, 2001, near Nadi, Fiji, and fell into the South Pacific Ocean.
On February 21, 2008, a disabled U.S. spy satellite, USA-193, was hit at an altitude of approximately 246 kilometers (153 mi) with an SM-3 missile fired from the U.S. Navy cruiser Lake Erie off the coast of Hawaii. The satellite was inoperative, having failed to reach its intended orbit when it was launched in 2006. Due to its rapidly deteriorating orbit it was destined for uncontrolled reentry within a month. U.S. Department of Defense expressed concern that the 1,000-pound (450 kg) fuel tank containing highly toxic hydrazine might survive reentry to reach the Earth's surface intact. Several governments including those of Russia, China, and Belarus protested the action as a thinly-veiled demonstration of US anti-satellite capabilities.[80] China had previously caused an international incident when it tested an anti-satellite missile in 2007.
Closeup of Gemini 2 heat shield
Cross section of Gemini 2 heat shield
Reentradas atmosféricas exitosas desde velocidades orbitales
Manned orbital reentry, by country/governmental entity
- China – Shenzhou
- Soviet Union/ Russia – Vostok, Voskhod, Soyuz
- United States – Mercury, Gemini, Apollo, Space Shuttle
Manned orbital reentry, by commercial entity
- SpaceX – Dragon 2
Unmanned orbital reentry, by country/governmental entity
- China
- European Space Agency[81]
- India / Indian Space Research Organisation
- Japan
- Soviet Union/ Russia
- United States
Unmanned orbital reentry, by commercial entity
- SpaceX – Dragon
Reentradas atmosféricas seleccionadas
This list includes some notable atmospheric entries in which the spacecraft was not intended to be recovered, but was destroyed in the atmosphere.
Spacecraft | Reentry year |
---|---|
Phobos-Grunt | 2012 |
ROSAT | 2011 |
UARS | 2011 |
Mir | 2001 |
Skylab | 1979 |
Ver también
- Van Allen radiation belt – Zone of energetic charged particles around the planet Earth
- Aerocapture – Orbital transfer maneuver
- Decelerated micrometeorites
- Ionization blackout
- Intercontinental ballistic missile – Ballistic missile with a range of more than 5,500 kilometres
- Lander (spacecraft) – Type of spacecraft
- Landing footprint
- List of reentering space debris – Wikipedia list article
- NASA reentry prototypes
- Skip reentry
- Space capsule – Type of spacecraft
- Space Shuttle thermal protection system – Space Shuttle heat shielding system
- Paper plane launched from space
notas y referencias
- ^ "ATO: Airship To Orbit" (PDF). JP Aerospace.
- ^ GROSS, F. (1965). "Buoyant Probes into the Venus Atmosphere". Unmanned Spacecraft Meeting 1965. American Institute of Aeronautics and Astronautics. doi:10.2514/6.1965-1407.
- ^ Goddard, Robert H. (Mar 1920). "Report Concerning Further Developments". The Smithsonian Institution Archives. Archived from the original on 26 June 2009. Retrieved 2009-06-29.
- ^ Boris Chertok, "Rockets and People", NASA History Series, 2006
- ^ Hansen, James R. (Jun 1987). "Chapter 12: Hypersonics and the Transition to Space". Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917–1958. The NASA History Series. sp-4305. United States Government Printing. ISBN 978-0-318-23455-7.
- ^ Allen, H. Julian; Eggers, A. J. Jr. (1958). "A Study of the Motion and Aerodynamic Heating of Ballistic Missiles Entering the Earth's Atmosphere at High Supersonic Speeds" (PDF). NACA Annual Report. NASA Technical Reports. 44.2 (NACA-TR-1381): 1125–1140. Archived from the original (PDF) on October 13, 2015.
- ^ http://www.nasa.gov/pdf/501326main_TA09-EDL-DRAFT-Nov2010-A.pdf
- ^ Graves, Claude A.; Harpold, Jon C. (March 1972). Apollo Experience Report - Mission Planning for Apollo Entry (PDF). NASA Technical Note (TN) D-6725.
The purpose of the Apollo entry maneuver is to dissipate the energy of a spacecraft traveling at high speed through the atmosphere of the earth so that the flight crew, their equipment, and their cargo are returned safely to a preselected location on the surface of the earth. This purpose must be accomplished while stresses on both the spacecraft and the flight crew are maintained within acceptable limits.
- ^ Przadka, W.; Miedzik, J.; Goujon-Durand, S.; Wesfreid, J.E. "The wake behind the sphere; analysis of vortices during transition from steadiness to unsteadiness" (PDF). Polish french cooperation in fluid research. Archive of Mechanics., 60, 6, pp. 467–474, Warszawa 2008. Received May 29, 2008; revised version November 13, 2008. Retrieved 3 April 2015.
- ^ a b Fay, J. A.; Riddell, F. R. (February 1958). "Theory of Stagnation Point Heat Transfer in Dissociated Air" (PDF). Journal of the Aeronautical Sciences. 25 (2): 73–85. doi:10.2514/8.7517. Archived from the original (PDF Reprint) on 2005-01-07. Retrieved 2009-06-29.
- ^ Hillje, Ernest R., "Entry Aerodynamics at Lunar Return Conditions Obtained from the Flight of Apollo 4 (AS-501)," NASA TN D-5399, (1969).
- ^ Whittington, Kurt Thomas. "A Tool to Extrapolate Thermal Reentry Atmosphere Parameters Along a Body in Trajectory Space" (PDF). NCSU Libraries Technical Reports Repository. A thesis submitted to the Graduate Faculty of North Carolina State University in partial fulfillment of the requirements for the degree of Master of Science Aerospace Engineering Raleigh, North Carolina 2011, pp.5. Retrieved 5 April 2015.
- ^ Regan, Frank J. and Anadakrishnan, Satya M., "Dynamics of Atmospheric Re-Entry", AIAA Education Series, American Institute of Aeronautics and Astronautics, Inc., New York, ISBN 1-56347-048-9, (1993).
- ^ a b Johnson, Sylvia M.; Squire, Thomas H.; Lawson, John W.; Gusman, Michael; Lau, K-H; Sanjuro, Angel (30 January 2014). Biologically-Derived Photonic Materials for Thermal Protection Systems (PDF). 38th Annual Conference on Composites, Materials, and Structures January 27–30, 2014.
- ^ Ionization And Dissociation Effects On Hypersonic Boundary-Layer Stability
- ^ "Equations, tables, and charts for compressible flow" (PDF). NACA Annual Report. NASA Technical Reports. 39 (NACA-TR-1135): 613–681. 1953.
- ^ Kenneth Iliff and Mary Shafer, Space Shuttle Hypersonic Aerodynamic and Aerothermodynamic Flight Research and the Comparison to Ground Test Results, Page 5-6
- ^ Lighthill, M.J. (Jan 1957). "Dynamics of a Dissociating Gas. Part I. Equilibrium Flow". Journal of Fluid Mechanics. 2 (1): 1–32. Bibcode:1957JFM.....2....1L. doi:10.1017/S0022112057000713.
- ^ Freeman, N.C. (Aug 1958). "Non-equilibrium Flow of an Ideal Dissociating Gas". Journal of Fluid Mechanics. 4 (4): 407–425. Bibcode:1958JFM.....4..407F. doi:10.1017/S0022112058000549.
- ^ Entry Aerodynamics at Lunar Return Conditions Obtained from the Fliigh of Apollo 4, Ernest R. Hillje, NASA, TN: D-5399, accessed 29 December 2018.
- ^ Overview of the Mars Sample Return Earth Entry Vehicle, NASA, accessed 29 December 2018.
- ^ Parker, John and C. Michael Hogan, "Techniques for Wind Tunnel assessment of Ablative Materials", NASA Ames Research Center, Technical Publication, August, 1965.
- ^ Hogan, C. Michael, Parker, John and Winkler, Ernest, of NASA Ames Research Center, "An Analytical Method for Obtaining the Thermogravimetric Kinetics of Char-forming Ablative Materials from Thermogravimetric Measurements", AIAA/ASME Seventh Structures and Materials Conference, April, 1966
- ^ "Arc Jet Complex". www.nasa.gov. NASA. Retrieved 2015-09-05.
- ^ Di Benedetto, A.T.; Nicolais, L.; Watanabe, R. (1992). Composite materials : proceedings of Symposium A4 on Composite Materials of the International Conference on Advanced Materials – ICAM 91, Strasbourg, France, 27–29 May 1991. Amsterdam: North-Holland. p. 111. ISBN 978-0444893567.
- ^ Tran, Huy; Michael Tauber; William Henline; Duoc Tran; Alan Cartledge; Frank Hui; Norm Zimmerman (1996). Ames Research Center Shear Tests of SLA-561V Heat Shield Material for Mars-Pathfinder (PDF) (Technical report). NASA Ames Research Center. NASA Technical Memorandum 110402.
- ^ Lachaud, Jean; N. Mansour, Nagi (June 2010). A pyrolysis and ablation toolbox based on OpenFOAM (PDF). 5th OpenFOAM Workshop. Gothenburg, Sweden. p. 1.
- ^ Tran, Huy K, et al., "Qualification of the forebody heat shield of the Stardust's Sample Return Capsule", AIAA, Thermophysics Conference, 32nd, Atlanta, GA; 23–25 June 1997.
- ^ "Stardust – Cool Facts". stardust.jpl.nasa.gov.
- ^ a b c Chambers, Andrew; Dan Rasky (2010-11-14). "NASA + SpaceX Work Together". NASA. Archived from the original on 2011-04-16. Retrieved 2011-02-16.
SpaceX undertook the design and manufacture of the reentry heat shield; it brought speed and efficiency that allowed the heat shield to be designed, developed, and qualified in less than four years.'
- ^ "SpaceX Manufactured Heat Shield Material Passes High Temperature Tests Simulating Reentry Heating Conditions of Dragon Spacecraft". www.spaceref.com.
- ^ Dragon could visit space station next, msnbc.com, 2010-12-08, accessed 2010-12-09.
- ^ Chaikin, Andrew (January 2012). "1 visionary + 3 launchers + 1,500 employees = ? : Is SpaceX changing the rocket equation?". Air & Space Smithsonian. Retrieved 2016-06-03.
SpaceX's material, called PICA-X, is 1/10th as expensive than the original [NASA PICA material and is better], ... a single PICA-X heat shield could withstand hundreds of returns from low Earth orbit; it can also handle the much higher energy reentries from the Moon or Mars.
- ^ NASA TV broadcast for the Crew Dragon Demo-2 mission departure from the ISS, NASA, 1 August 2020.
- ^ Tran, Huy K., et al., "Silicone impregnated reusable ceramic ablators for Mars follow-on missions," AIAA-1996-1819, Thermophysics Conference, 31st, New Orleans, June 17–20, 1996.
- ^ Flight-Test Analysis Of Apollo Heat-Shield Material Using The Pacemaker Vehicle System NASA Technical Note D-4713, pp. 8, 1968–08, accessed 2010-12-26. "Avcoat 5026-39/HC-G is an epoxy novolac resin with special additives in a fiberglass honeycomb matrix. In fabrication, the empty honeycomb is bonded to the primary structure and the resin is gunned into each cell individually. ... The overall density of the material is 32 lb/ft3 (512 kg/m3). The char of the material is composed mainly of silica and carbon. It is necessary to know the amounts of each in the char because in the ablation analysis the silica is considered to be inert, but the carbon is considered to enter into exothermic reactions with oxygen. ... At 2160O R (12000 K), 54 percent by weight of the virgin material has volatilized and 46 percent has remained as char. ... In the virgin material, 25 percent by weight is silica, and since the silica is considered to be inert the char-layer composition becomes 6.7 lb/ft3 (107.4 kg/m3) of carbon and 8 lb/ft3 (128.1 kg/m3) of silica."
- ^ NASA.gov NASA Selects Material for Orion Spacecraft Heat Shield, 2009-04-07, accessed 2011-01-02.
- ^ Flightglobal.com NASA's Orion heat shield decision expected this month 2009-10-03, accessed 2011-01-02
- ^ "Company Watch – NASA. – Free Online Library". www.thefreelibrary.com.
- ^ Shao, Gaofeng; et al. (2019). "Improved oxidation resistance of high emissivity coatings on fibrous ceramic for reusable space systems". Corrosion Science. 146: 233–246. arXiv:1902.03943. doi:10.1016/j.corsci.2018.11.006. S2CID 118927116.
- ^ Columbia Accident Investigation Board report
- ^ Shuttle Evolutionary History
- ^ X-33 Heat Shield Development report
- ^ "Archived copy" (PDF). Archived from the original (PDF) on 2005-12-15. Retrieved 2006-04-09.CS1 maint: archived copy as title (link)
- ^ sharp structure homepage w left Archived October 16, 2015, at the Wayback Machine
- ^ Why Elon Musk Turned to Stainless Steel for SpaceX's Starship Mars Rocket, Mike Wall, space.com, 23 January 2019, accessed 23 March 2019.
- ^ SpaceX CEO Elon Musk explains Starship's "transpiring" steel heat shield in Q&A, Eric Ralph, Teslarati News, 23 January 2019, accessed 23 March 2019
- ^ "- J2T-200K & J2T-250K".
- ^ "How SpaceShipOne Works". 20 June 2004.
- ^ Chapman, Dean R. (May 1958). "An approximate analytical method for studying reentry into planetary atmospheres" (PDF). NACA Technical Note 4276: 38. Archived from the original (PDF) on 2011-04-07.
- ^ a b c NASA Launches New Technology: An Inflatable Heat Shield, NASA Mission News, 2009-08-17, accessed 2011-01-02.
- ^ "Inflatable Re-Entry Technologies: Flight Demonstration and Future Prospects" (PDF).
- ^ Inflatable Reentry and Descent Technology (IRDT) Archived 2015-12-31 at the Wayback Machine Factsheet, ESA, September, 2005
- ^ IRDT demonstration missions Archived 2016-12-07 at the Wayback Machine
- ^ Hughes, Stephen J. "Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology Development Overview" (PDF). www.nasa.gov. NASA. Archived from the original (PDF) on 26 January 2017. Retrieved 28 March 2017.
- ^ Cheatwood, Neil (29 June 2016). "Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology" (PDF). www.nasa.gov. NASA. Retrieved 28 March 2017.
- ^ Launch Vehicle Recovery and Reuse
- ^ NOAA finalizes secondary payload for JPSS-2 launch
- ^ a b c Pavlosky, James E., St. Leger, Leslie G., "Apollo Experience Report - Thermal Protection Subsystem," NASA TN D-7564, (1974).
- ^ William Harwood (2008). "Whitson describes rough Soyuz entry and landing". Spaceflight Now. Retrieved July 12, 2008.
- ^ Spacecraft Reentry FAQ: How much material from a satellite will survive reentry? Archived March 2, 2014, at the Wayback Machine
- ^ NASA - Frequently Asked Questions: Orbital Debris Archived March 11, 2014, at the Wayback Machine
- ^ "Animation52-desktop". www.aerospace.org. Archived from the original on 2014-03-02. Retrieved 2013-03-04.
- ^ "3-2-2-1 Settlement of Claim between Canada and the Union of Soviet Socialist Republics for Damage Caused by "Cosmos 954" (Released on April 2, 1981)". www.jaxa.jp.
- ^ Hanslmeier, Arnold (2002). The sun and space weather. Dordrecht; Boston: Kluwer Academic Publishers. p. 269. ISBN 9781402056048.
- ^ Lamprecht, Jan (1998). Hollow planets : a feasibility study of possible hollow worlds. Austin, Texas: World Wide Pub. p. 326. ISBN 9780620219631.
- ^ Elkins-Tanton, Linda (2006). The Sun, Mercury, and Venus. New York: Chelsea House. p. 56. ISBN 9780816051939.
- ^ aero.org, Spacecraft Reentry FAQ: Archived May 13, 2012, at the Wayback Machine
- ^ Astronautix, Salyut 7.
- ^ "Salyut 7, Soviet Station in Space, Falls to Earth After 9-Year Orbit" New York Times
- ^ David, Leonard (7 September 2011). "Huge Defunct Satellite to Plunge to Earth Soon, NASA Says". Space.com. Retrieved 10 September 2011.
- ^ "Final Update: NASA's UARS Re-enters Earth's Atmosphere". Retrieved 2011-09-27.
- ^ "aerospace.org Tiangong-1 Reentry". Archived from the original on 2018-04-04. Retrieved 2018-04-02.
- ^ Jones, Morris (30 March 2016). "Has Tiangong 1 gone rogue". Space Daily. Retrieved 22 September 2016.
- ^ 18 Space Control Squadron [@18SPCS] (11 May 2020). "#18SPCS has confirmed the reentry of the CZ-5B R/B (#45601, 2020-027C) at 08:33 PDT on 11 May, over the Atlantic Ocean. The #CZ5B launched China's test crew capsule on 5 May 2020. #spaceflightsafety" (Tweet). Retrieved 11 May 2020 – via Twitter.
- ^ Clark, Stephen. "China's massive Long March 5B's rocket falls out of orbit over Atlantic Ocean – Spaceflight Now". Retrieved 2020-05-12.
- ^ "Bridenstine Criticizes Uncontrolled Long March 5B Stage Reentry – Parabolic Arc". Retrieved 2020-05-16.
- ^ O'Callaghan, Jonathan. "Chinese Rocket Debris May Have Fallen On Several African Villages After An Uncontrolled Re-Entry". Forbes. Retrieved 2020-05-13.
- ^ Fernando, Benjamin; Wojcicka, Natalia; Froment, Marouchka; Maguire, Ross; Staehler, Simon; Rolland, Lucie; Collins, Gareth; Karatekin, Ozgur; Larmat, Carene; Sansom, Eleanor; Teanby, Nicholas (2020-12-02). "Listening for the Landing: Detecting Perseverance's landing with InSight". Cite journal requires
|journal=
(help) - ^ Gray, Andrew (2008-02-21). "U.S. has high confidence it hit satellite fuel tank". Reuters. Archived from the original on 25 February 2008. Retrieved 2008-02-23.
- ^ "IXV flight profile". European Space Agency.
Otras lecturas
- Launius, Roger D.; Jenkins, Dennis R. (October 10, 2012). Coming Home: Reentry and Recovery from Space. NASA. ISBN 9780160910647. OCLC 802182873. Retrieved August 21, 2014.
- Martin, John J. (1966). Atmospheric Entry – An Introduction to Its Science and Engineering. Old Tappan, New Jersey: Prentice-Hall.
- Regan, Frank J. (1984). Re-Entry Vehicle Dynamics (AIAA Education Series). New York: American Institute of Aeronautics and Astronautics, Inc. ISBN 978-0-915928-78-1.
- Etkin, Bernard (1972). Dynamics of Atmospheric Flight. New York: John Wiley & Sons, Inc. ISBN 978-0-471-24620-6.
- Vincenti, Walter G.; Kruger Jr, Charles H. (1986). Introduction to Physical Gas Dynamics. Malabar, Florida: Robert E. Krieger Publishing Co. ISBN 978-0-88275-309-6.
- Hansen, C. Frederick (1976). Molecular Physics of Equilibrium Gases, A Handbook for Engineers. NASA. Bibcode:1976mpeg.book.....H. NASA SP-3096.
- Hayes, Wallace D.; Probstein, Ronald F. (1959). Hypersonic Flow Theory. New York and London: Academic Press. A revised version of this classic text has been reissued as an inexpensive paperback: Hayes, Wallace D. (1966). Hypersonic Inviscid Flow. Mineola, New York: Dover Publications. ISBN 978-0-486-43281-6. reissued in 2004
- Anderson, John D. Jr. (1989). Hypersonic and High Temperature Gas Dynamics. New York: McGraw-Hill, Inc. ISBN 978-0-07-001671-2.
enlaces externos
- Aerocapture Mission Analysis Tool (AMAT) provides preliminary mission analysis and simulation capability for atmospheric entry vehicles at various Solar System destinations.
- Center for Orbital and Reentry Debris Studies (The Aerospace Corporation)
- Apollo Atmospheric Entry Phase, 1968, NASA Mission Planning and Analysis Division, Project Apollo. video (25:14).
- Buran's heat shield
- Encyclopedia Astronautica article on the history of space rescue crafts, including some reentry craft designs.