再入：修订间差异
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|T=zh-hans:再入; zh-hant:重返; |
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|1=zh-hans:再入; zh-hant:重返; |
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[[File:Entry.jpg|thumb|300px|[[火星探测漫游者]]（MER）的[[整流罩]]在进入火星大气时情形的艺术想象图]] |
[[File:Entry.jpg|thumb|300px|[[火星探测漫游者]]（MER）的[[整流罩]]在进入火星大气时情形的艺术想象图]] |
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'''重返'''（{{lang-en|Re-entry}}，-{zh-hans:正; zh-hant:簡;}-體中文譯為'''-{zh-hans:重返; zh-hant:再入;}-'''）是指自然物体或人造物体从[[外層空间]]进入[[行星]][[大气层]]的运动过程。如果人造物体（如[[人造卫星]]、[[飞船]]、[[火箭]][[导弹]]、[[空天飞机]]等）离开[[地球大气层]]，再从[[外太空]]重新进入[[地球]]大气层的运动，称为“'''重返'''”（reentry）大气层。<ref>{{cite web |title=re-entry - 重返 |url=https://terms.naer.edu.tw/detail/952464/ |website=terms.naer.edu.tw |access-date=2021-12-13 |archive-date=2021-12-13 |archive-url=https://web.archive.org/web/20211213142001/https://terms.naer.edu.tw/detail/952464/ }}</ref>例如，[[流星]]進入地球大氣墜落的[[隕石]]不算重返。 |
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弹道弹头和消耗性飞行器在再入大气层时不需要减速，事实上，它们被制成[[流线型]]以保持速度。此外，从[[临近空间]]慢速返回地球（如从[[气球]]上[[跳伞]]）不需要热屏蔽，因为[[重力加速度]]不会对从大气层本身（或在大气层上方不远处）开始[[相對靜止|相对静止]]的物体产生足够的速度以造成显著的加热。 |
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例如，火星探测器在火星表面软着陆，就只能称为“进入”火星大气层，而不能称“再入”火星大气层。 |
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{{Citation needed|对于地球来说，重返发生在[[卡門線|卡门线]]处，[[海拔]]100公里（62英里，54海里）上方的高度，而在[[金星]]则发生在250公里（160英里，30海里）处，[[火星]]在约80公里（50英里，43海里）处。|time=2023-06-12}} |
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| ⚫ | 人造或者自然物体从[[宇宙空间]]进入[[天体]]的[[大气层]]的过程被称作进入大气层（Atmospheric entry），在[[地球]]的场合指的是从宇宙空间一侧越过[[海拔]]为100km的[[卡门线]]的过程。{{better source|reason= 引用[2]没有解释为何再入发生于越过卡门线后。|date=2023年06月}}从地面发射后离开大气层的人造航天载具重新进入大气层的过程被称作'''重返大气层'''（Atmospheric reentry）或'''重返'''（Re-entry）。<ref>{{cite web |title=atmospheric entry - 進入大氣 |url=https://terms.naer.edu.tw/detail/368639/?index=1 |website=terms.naer.edu.tw |accessdate=2021-12-13 |archive-date=2021-12-13 |archive-url=https://web.archive.org/web/20211213142006/https://terms.naer.edu.tw/detail/368639/?index=1 }}</ref> |
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返回大气层根据其目的和过程被分为以下类型： |
返回大气层根据其目的和过程被分为以下类型： |
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| 把[[航天器]]安全降落到行星表面 || 是 || 否 |
| 把[[航天器]]安全降落到行星表面 || 是 || 否 |
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| [[洲际弹道导弹]]的[[弹道飞行]]后半程 || 是 || |
| [[洲际弹道导弹]]的[[弹道飞行]]后半程 || 是 || 是 |
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| 人为消灭航天器或[[太空垃圾]] || 否，或任由其轨道自然衰减 || 是 |
| 人为消灭航天器或[[太空垃圾]] || 否，或任由其轨道自然衰减 || 是 |
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==歷史== |
==歷史== |
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這種雙層隔熱板概念在1920年由[[羅伯特·戈達德]]提出，他說："[[流星]]進入大氣層的速度高達每秒30英里，但內部依然寒冷。因此，假如再返回物的表面覆蓋一層抗高溫（不易變質及難熔解）的物質後再用一層不太會導熱的耐高溫物質，這樣物體表面就不會受到太多的侵蝕"（節錄） |
{{Citation needed|這種雙層隔熱板概念在1920年由[[羅伯特·戈達德]]提出，他說："[[流星]]進入大氣層的速度高達每秒30英里，但內部依然寒冷。因此，假如再返回物的表面覆蓋一層抗高溫（不易變質及難熔解）的物質後再用一層不太會導熱的耐高溫物質，這樣物體表面就不會受到太多的侵蝕"（節錄）。|date=2023-06-12}} |
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而第一次實際應用到此系統是在洲际彈道導彈的 |
而第一次實際應用到此系統是在洲际彈道導彈的重返速度增加所導致的气动力加热。早期的彈道導彈，如[[V2火箭]]，並沒有此問題。而中程彈道導彈，如蘇聯的[[R-5]]（有1200km）的射程，就需要陶瓷複合材料來保護。首個[[洲際彈道導彈]]（ICBM，射程達8000至12000km），則已正式進入了現代保護材料的時代。在美國，這技術是由{{tsl|en|Ames Research Center|艾姆斯研究中心（NASA Ames）}}任職的{{tsl|en|Harry Julian Allen|哈立·朱利安·艾倫}}率先開發。而蘇聯的Yuri A. Dunaev也曾在列寧格勒物理技術研究所開發類似的技術。 |
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==飛行器的形狀== |
==飛行器的形狀== |
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===鈍形飛行器=== |
===鈍形飛行器=== |
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美國國家航空諮詢委員會的{{tsl|en|Harry Julian Allen| |
[[美國國家航空諮詢委員會]]的{{tsl|en|Harry Julian Allen|哈立·朱利安·艾倫}}、{{tsl|en|Alfred J. Eggers|阿爾弗雷德·J·艾格斯}}在1951年發現了鈍形（高阻）[[防热盾|防热大底]]具有最佳效果。因為返回式航天器的气动加熱與[[阻力係數]]成反比，即阻力愈大，熱負荷愈低。钝形重返舱使得气体不能快速离开，成为气垫层隔开了冲击波与加热振动层，使得大部分热空气不再直接接触重返舱，热能保持在冲击波气体中并在大气层中扩散。 |
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艾倫和 |
艾倫和艾格斯的發現，最初被視為軍事秘密，但於1958年出版。鈍形理論的設計成為可行的隔熱板，都體現在水星、雙子星和阿波羅太空艙，使宇航員返回火熱的地球大氣層時仍生存。蘇聯的[[R-7]]洲際彈道導彈於1957年使用尖鼻的彈頭成功首次試射，但擊中目標區10公里以外，因而改為鈍鼻的彈頭。蘇聯的隔熱層由多層玻璃纖維與石棉textolite組成。 |
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== |
==重返飞行器形状== |
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===球形或球形部分=== |
===球形或球形部分=== |
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[[File: |
[[File:Apollo Block1 CM.PNG|thumb|阿波罗飞船指令舱的重返形状]] |
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| ⚫ | 1950年代到1960年代，易于从理论上用Fay-Riddell方程建模分析。<ref name="Fay-Riddell">{{cite journal |last1=Fay |first1=J. A. |last2=Riddell |first2=F. R. |title=Theory of Stagnation Point Heat Transfer in Dissociated Air |journal=Journal of the Aeronautical Sciences |volume=25 |pages=73–85 |date=February 1958 |url=http://pdf.aiaa.org/downloads/TOCPDFs/36_373-386.pdf |format=PDF Reprint |accessdate=2009-06-29 |issue=2 |doi=10.2514/8.7517 |deadurl=yes |archiveurl=https://web.archive.org/web/20050107202757/http://pdf.aiaa.org/downloads/TOCPDFs/36_373-386.pdf |archivedate=2005-01-07 }}</ref> 当时没有高速计算机，高速空气动力学还处于萌芽阶段。苏联的[[東方計劃|东方飞船]]、[[上升计划|上升飞船]]、以及火星、金星探测器的下降舱；美国的{{tsl|en|Apollo Command/Service Module|阿波罗飞船指令舱}}都采用了球形防热大底。阿波罗飞船重返时的攻角−27°，升阻比0.368<ref>{{Cite web |url=http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19690029435_1969029435.pdf |title=Hillje, Ernest R., "Entry Aerodynamics at Lunar Return Conditions Obtained from the Flight of Apollo 4 (AS-501)," NASA TN D-5399, (1969). |accessdate=2014-11-07 |archive-date=2020-11-11 |archive-url=https://web.archive.org/web/20201111215043/http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19690029435_1969029435.pdf |dead-url=no }}</ref>。[[联盟计划|联盟飞船]]、[[探測器號系列探測器|月球取样返回探测器]]、[[双子座飞船]]、[[水星计划|水星号飞船]]都是如此设计。即使这少量的升力也使得从弹道式重返8-9g的峰值加速度减小到4-5g，同时大大减少了峰值气动加热。 |
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[[File:Apollo Block1 CM.PNG|thumb|阿波罗飞船指令舱的再入形状]] |
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| ⚫ | 1950年代到1960年代，易于从理论上用Fay-Riddell方程建模分析。<ref name="Fay-Riddell">{{cite journal |last1=Fay |first1=J. A. |last2=Riddell |first2=F. R. |title=Theory of Stagnation Point Heat Transfer in Dissociated Air |journal=Journal of the Aeronautical Sciences |volume=25 |pages=73–85 |date=February 1958 |url=http://pdf.aiaa.org/downloads/TOCPDFs/36_373-386.pdf |format=PDF Reprint |accessdate=2009-06-29 |issue=2 |doi=10.2514/8.7517}}</ref> 当时没有高速计算机，高速空气动力学还处于萌芽阶段。苏联的[[東方計劃|东方飞船]]、[[上升计划|上升飞船]]、以及火星、金星探测器的下降舱；美国的{{tsl|en|Apollo Command/Service Module|阿波罗飞船指令舱}}都采用了球形防热大底。阿波罗飞船 |
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===球-圆锥形=== |
===球-圆锥形=== |
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[[Image:Galileo probe.jpg|thumb|left|最终组装时的伽利略探测器]] |
[[Image:Galileo probe.jpg|thumb|left|最终组装时的伽利略探测器]] |
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球-圆锥形是指截头圆锥与球形部分的结合。这具有更好的动态稳定性。 |
球-圆锥形是指截头圆锥与球形部分的结合。这具有更好的动态稳定性。 |
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[[Image:Missile Maintainer inspects missile guidance system of the LGM-30G Minuteman ICBM.jpg|thumb|upright|LGM-30民兵洲际导弹的 |
[[Image:Missile Maintainer inspects missile guidance system of the LGM-30G Minuteman ICBM.jpg|thumb|upright|LGM-30民兵洲际导弹的重返段]] |
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美国最早的该构形的 |
美国最早的该构形的重返舱是[[通用电气]]于1955年开发的Mk-2 RV。使用了基于金属防热大底的辐射热防护系统(TPS)。Mk-2作为武器投射系统具有很大缺陷，由于低{{tsl|en|ballistic coefficient|弹道系数}}使得其在上层大气中飞行时间太长，产生一股金属蒸汽尾流，极易于被雷达发现。 |
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[[Image:Mk 6 reentry vehicle on display at National Atomic Museum.jpg|thumb|upright|left|Mk-6 |
[[Image:Mk 6 reentry vehicle on display at National Atomic Museum.jpg|thumb|upright|left|Mk-6 重返舱]] |
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通用电气研制的下一代 |
通用电气研制的下一代重返舱是Mk-6，采用尼龙酚醛材质的防热大底，其效果非常好以至于可以大大减小锥体半角到12.5°. Mk-6的重返质量3360 kg，长3.1m。随着核武器小型化与烧蚀材料的进展，重返舱变得更轻、锥体半角减小到10°-11°. |
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[[Image:Rv film pod.jpg|thumb|upright|"发现者"侦察卫星的胶片返回舱]] |
[[Image:Rv film pod.jpg|thumb|upright|"发现者"侦察卫星的胶片返回舱]] |
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美国 |
美国[[日冕侦察卫星]]是第一种非武器战斗部的重返舱。1959年2月28日首次发射。 February 1959). |
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[[File:Viking_spacecraft.jpg|thumb|upright|携带着陆器的海盗号轨道器]] |
[[File:Viking_spacecraft.jpg|thumb|upright|携带着陆器的海盗号轨道器]] |
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不同于军事目的的返回舱，空间探测器的返回舱采取更大的锥体半角显然可以减少对烧蚀材料的需要，降低死重。[[伽利略号探测器]]的下降舱的锥体半角达到了45°，[[海盗号]]火星着陆舱的锥体半角达到70° |
不同于军事目的的返回舱，空间探测器的[[返回舱]]采取更大的锥体半角显然可以减少对烧蚀材料的需要，{{Citation needed|降低死重|time=2021-12-13}}。[[伽利略号探测器]]的下降舱的锥体半角达到了45°，[[海盗号]]火星着陆舱的锥体半角达到70° |
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[[File:Stardust Capsule on Ground 011506.JPG|thumb|left|[[星尘号]]彗星采样 |
[[File:Stardust Capsule on Ground 011506.JPG|thumb|left|[[星尘号]]彗星采样重返舱]] |
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===双锥形=== |
===双锥形=== |
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双锥形是在球形-锥形上又增加了一个截头锥形。这具有非常好的升阻比，达到1.0以上。 |
双锥形是在球形-锥形上又增加了一个截头锥形。这具有非常好的升阻比，达到1.0以上。 |
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===非轴对称形状=== |
===非轴对称形状=== |
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用于载人 |
用于载人重返舱，如采用三角翼的[[航天飞机]]、[[暴风雪号航天飞机]]，以及[[升力体]]如{{tsl|en|Martin X-23A PRIME|X-23 PRIME}}。 |
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==激波层气体物理== |
==激波层气体物理== |
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防热大底设计的经验规则是：[[激波层]]气体峰值温度（[[开尔文]]为单位）等于进入速度（单位m/s）。例如，宇宙飞船以7.8 km/s进入大气层，激波层气体峰值温度7800 K。因为[[动能]]增加与速度的平方成正比，而气体的[[比热容]]随温度而大幅度增加（这不同于固体在通常条件下可以假设比热容不变）。 |
防热大底设计的经验规则是：[[激波层]]气体峰值温度（[[开尔文]]为单位）等于进入速度（单位m/s）。例如，宇宙飞船以7.8 km/s进入大气层，激波层气体峰值温度7800 K。因为[[动能]]增加与速度的平方成正比，而气体的[[比热容]]随温度而大幅度增加（这不同于固体在通常条件下可以假设比热容不变）。 |
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在典型的 |
在典型的重返温度，激波层大气是被[[电离]]与[[解离]]的。这种化学解离必须一些物理模型以描述激波层的热与化学性质。对于设计防热大底的航宇工程师有4种气体基本物理模型： |
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===理想气体模型=== |
===理想气体模型=== |
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几乎所有的航 |
几乎所有的航太工程师在本科时学过[[理想气体|理想气体模型]]。大部分的理想气体方程与对应的表与图在NACA Report 1135中给出。<ref>{{cite journal |title=Equations, tables, and charts for compressible flow |publisher=NASA Technical Reports |issue=NACA-TR-1135 |year=1953 |url=http://ntrs.nasa.gov/search.jsp?R=451364&id=3qs=Ntt%3D1135%26Ntk%3Dall%26Ntx%3Dmode%2Bmatchall%26Ns%3DPublicationYear%257c1%26N%3D17 |journal=NACA Annual Report |volume=39 |pages=611–681 |access-date=2014-11-08 |archive-date=2020-09-17 |archive-url=https://web.archive.org/web/20200917170551/https://ntrs.nasa.gov/search.jsp?R=451364&id=3qs=Ntt%3D1135%26Ntk%3Dall%26Ntx%3Dmode%2Bmatchall%26Ns%3DPublicationYear%257c1%26N%3D17 |dead-url=no }}</ref> NACA Report 1135的摘要经常出现在热动力学课本的附录被航宇工程师熟悉。 |
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理想气体理论非常精巧并在设计飞机时非常有用，但它假定气体是化学惰性的。从飞机设计角度，大气温度低于550 K时可以假定气体为惰性的。气温550 K时理想气体理论开始出现问题而气温超过2000 K将不再适用，这时防热大底设计者必须使用 |
理想气体理论非常精巧并在设计飞机时非常有用，但它假定气体是化学惰性的。从飞机设计角度，大气温度低于550 K时可以假定气体为惰性的。气温550 K时理想气体理论开始出现问题而气温超过2000 K将不再适用，这时防热大底设计者必须使用真实气体模型''。'' |
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== 再入角 == |
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===Real (equilibrium) gas model=== |
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再入角的大小直接影响到航天器在大气层里的航程。若再入角过小，航天器可能只在大气层边缘掠过而无法进入；若过大则会受到较大的空气阻力，可能会使气动力加热严重进而导致航天器燃烧。一般无人航天器再入大气层时，再入角在3°~8°之间，不会超过10°。<ref>{{Cite journal |author=赵玉晖 |author2=侯锡云 |author3=刘林 |title=月球返回轨道再入角变化特征 |journal=飞行器测控学报 |issue=2010,29(05) |page=75-79 |via=中国知网}}</ref>为使最大制动过载不超过人体所能耐受的限度（10倍重力加速度），以[[第一宇宙速度]]再入的载人飞船必须以小于3°的再入角进入大气层。 |
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An entry vehicle's pitching moment can be significantly influenced by real-gas effects. Both the Apollo-CM and the Space Shuttle were designed using incorrect pitching moments determined through inaccurate real-gas modeling. The Apollo-CM's trim-angle angle of attack was higher than originally estimated, resulting in a narrower lunar return entry corridor. The actual aerodynamic center of the [[Space Shuttle Columbia|''Columbia'']] was upstream from the calculated value due to real-gas effects. On ''Columbia''’s maiden flight ([[STS-1]]), astronauts [[John Young (astronaut)|John W. Young]] and [[Robert Crippen]] had some anxious moments during reentry when there was concern about losing control of the vehicle.<ref>Kenneth Iliff and Mary Shafer, ''Space Shuttle Hypersonic Aerodynamic and Aerothermodynamic Flight Research and the Comparison to Ground Test Results'', Page 5-6</ref> |
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An equilibrium real-gas model assumes that a gas is chemically reactive, but also assumes all chemical reactions have had time to complete and all components of the gas have the same temperature (this is called ''[[thermodynamic equilibrium]]''). When air is processed by a shock wave, it is superheated by compression and chemically dissociates through many different reactions. Direct friction upon the reentry object is not the main cause of shock-layer heating. It is caused mainly from [[isentropic process|isentropic]] heating of the air molecules within the compression wave. Friction based entropy increases of the molecules within the wave also account for some heating.{{or|date=November 2013}} The distance from the shock wave to the [[stagnation point]] on the entry vehicle's leading edge is called ''shock wave stand off''. An approximate rule of thumb for shock wave standoff distance is 0.14 times the nose radius. One can estimate the time of travel for a gas molecule from the shock wave to the stagnation point by assuming a free stream velocity of 7.8 km/s and a nose radius of 1 meter, i.e., time of travel is about 18 microseconds. This is roughly the time required for shock-wave-initiated chemical dissociation to approach [[chemical equilibrium]] in a shock layer for a 7.8 km/s entry into air during peak heat flux. Consequently, as air approaches the entry vehicle's stagnation point, the air effectively reaches chemical equilibrium thus enabling an equilibrium model to be usable. For this case, most of the shock layer between the shock wave and leading edge of an entry vehicle is chemically reacting and ''not'' in a state of equilibrium. The [[wikt:Appendix:Glossary of atmospheric reentry#F|Fay-Riddell equation]],<ref name="Fay-Riddell"/> which is of extreme importance towards modeling heat flux, owes its validity to the stagnation point being in chemical equilibrium. The time required for the shock layer gas to reach equilibrium is strongly dependent upon the shock layer's pressure. For example, in the case of the [[Galileo (spacecraft)|Galileo Probe]]'s entry into Jupiter's atmosphere, the shock layer was mostly in equilibrium during peak heat flux due to the very high pressures experienced (this is counterintuitive given the free stream velocity was 39 km/s during peak heat flux). |
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==外部鏈結== |
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Determining the thermodynamic state of the stagnation point is more difficult under an equilibrium gas model than a perfect gas model. Under a perfect gas model, the ''ratio of specific heats'' (also called "isentropic exponent", [[adiabatic index]], "gamma" or "kappa") is assumed to be constant along with the [[gas constant]]. For a real gas, the ratio of specific heats can wildly oscillate as a function of temperature. Under a perfect gas model there is an elegant set of equations for determining thermodynamic state along a constant entropy stream line called the ''isentropic chain''. For a real gas, the isentropic chain is unusable and a ''Mollier diagram'' would be used instead for manual calculation. However, graphical solution with a Mollier diagram is now considered obsolete with modern heat shield designers using computer programs based upon a digital lookup table (another form of Mollier diagram) or a chemistry based thermodynamics program. The chemical composition of a gas in equilibrium with fixed pressure and temperature can be determined through the ''Gibbs free energy method''. [[Gibbs free energy]] is simply the total [[enthalpy]] of the gas minus its total [[entropy]] times temperature. A chemical equilibrium program normally does not require chemical formulas or reaction-rate equations. The program works by preserving the original elemental abundances specified for the gas and varying the different molecular combinations of the elements through numerical iteration until the lowest possible Gibbs free energy is calculated (a [[Newton's method|Newton-Raphson method]] is the usual numerical scheme). The data base for a Gibbs free energy program comes from spectroscopic data used in defining [[Partition function (statistical mechanics)|partition functions]]. Among the best equilibrium codes in existence is the program ''Chemical Equilibrium with Applications'' (CEA) which was written by Bonnie J. McBride and Sanford Gordon at NASA Lewis (now renamed "NASA Glenn Research Center"). Other names for CEA are the "Gordon and McBride Code" and the "Lewis Code". CEA is quite accurate up to 10,000 K for planetary atmospheric gases, but unusable beyond 20,000 K ([[double ionization]] is not modeled). [http://www.grc.nasa.gov/WWW/CEAWeb/ CEA can be downloaded from the Internet] along with full documentation and will compile on Linux under the G77 Fortran compiler. |
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*[https://www.nasa.gov/ames 艾姆斯研究中心（NASA Ames）] {{Wayback|url=https://www.nasa.gov/ames |date=20220515183601 }} |
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===Real (non-equilibrium) gas model=== |
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A non-equilibrium real gas model is the most accurate model of a shock layer's gas physics, but is more difficult to solve than an equilibrium model. The simplest non-equilibrium model is the ''Lighthill-Freeman model''.<ref>{{cite journal |last=Lighthill |first=M.J. |title=Dynamics of a Dissociating Gas. Part I. Equilibrium Flow |journal=Journal of Fluid Mechanics |volume=2 |pages=1–32 |date=Jan 1957 |doi=10.1017/S0022112057000713 |issue=1|bibcode = 1957JFM.....2....1L }}</ref><ref>{{cite journal |last=Freeman |first=N.C. |title=Non-equilibrium Flow of an Ideal Dissociating Gas |journal=Journal of Fluid Mechanics |volume=4 |pages=407–425 |date=Aug 1958 |doi=10.1017/S0022112058000549 |issue=4|bibcode = 1958JFM.....4..407F }}</ref> The Lighthill-Freeman model initially assumes a gas made up of a single diatomic species susceptible to only one chemical formula and its reverse; e.g., N<sub>2</sub> → N + N and N + N → N<sub>2</sub> (dissociation and recombination). Because of its simplicity, the Lighthill-Freeman model is a useful pedagogical tool, but is unfortunately too simple for modeling non-equilibrium air. Air is typically assumed to have a mole fraction composition of 0.7812 molecular nitrogen, 0.2095 molecular oxygen and 0.0093 argon. The simplest real gas model for air is the ''five species model'' which is based upon N<sub>2</sub>, O<sub>2</sub>, NO, N and O. The five species model assumes no ionization and ignores trace species like carbon dioxide. |
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When running a Gibbs free energy equilibrium program, the iterative process from the originally specified molecular composition to the final calculated equilibrium composition is essentially random and not time accurate. With a non-equilibrium program, the computation process is time accurate and follows a solution path dictated by chemical and reaction rate formulas. The five species model has 17 chemical formulas (34 when counting reverse formulas). The Lighthill-Freeman model is based upon a single ordinary differential equation and one algebraic equation. The five species model is based upon 5 ordinary differential equations and 17 algebraic equations. Because the 5 ordinary differential equations are loosely coupled, the system is numerically "stiff" and difficult to solve. The five species model is only usable for entry from [[low Earth orbit]] where entry velocity is approximately 7.8 km/s. For lunar return entry of 11 km/s, the shock layer contains a significant amount of ionized nitrogen and oxygen. The five species model is no longer accurate and a twelve species model must be used instead. High speed Mars entry which involves a carbon dioxide, nitrogen and argon atmosphere is even more complex requiring a 19 species model. |
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An important aspect of modeling 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 unsymmetric diatomic molecules; e.g., [[cyanogen]] (CN), [[carbon monoxide]], [[nitric oxide]] (NO), single ionized molecular nitrogen, et cetera. 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 [[quantum vibration|vibrational energy]] into radiant energy; i.e., radiative heat flux. The whole process takes place in less than a millisecond which makes modeling 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. |
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===Frozen gas model=== |
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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. |
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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]], et cetera) 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 modeling the flow in the wake of an entry vehicle is very difficult. [[Space Shuttle thermal protection system|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. |
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[[Category:太空探索]] |
[[Category:太空探索]] |
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[[Category:进入大气层]] |
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2023年11月8日 (三) 18:34的最新版本
重返（英語：Re-entry，正體中文譯為重返）是指自然物体或人造物体从外層空间进入行星大气层的运动过程。如果人造物体（如人造卫星、飞船、火箭导弹、空天飞机等）离开地球大气层，再从外太空重新进入地球大气层的运动，称为“重返”（reentry）大气层。[1]例如，流星進入地球大氣墜落的隕石不算重返。
弹道弹头和消耗性飞行器在再入大气层时不需要减速，事实上，它们被制成流线型以保持速度。此外，从临近空间慢速返回地球（如从气球上跳伞）不需要热屏蔽，因为重力加速度不会对从大气层本身（或在大气层上方不远处）开始相对静止的物体产生足够的速度以造成显著的加热。
对于地球来说，重返发生在卡门线处，海拔100公里（62英里，54海里）上方的高度，而在金星则发生在250公里（160英里，30海里）处，火星在约80公里（50英里，43海里）处。[來源請求]
原理
[编辑]人造或者自然物体从宇宙空间进入天体的大气层的过程被称作进入大气层（Atmospheric entry），在地球的场合指的是从宇宙空间一侧越过海拔为100km的卡门线的过程。[需要較佳来源]从地面发射后离开大气层的人造航天载具重新进入大气层的过程被称作重返大气层（Atmospheric reentry）或重返（Re-entry）。[2]
返回大气层根据其目的和过程被分为以下类型：
| 目的（原因） | 过程是否可控 | 是否破坏性 |
|---|---|---|
| 把航天器安全降落到行星表面 | 是 | 否 |
| 洲际弹道导弹的弹道飞行后半程 | 是 | 是 |
| 人为消灭航天器或太空垃圾 | 否，或任由其轨道自然衰减 | 是 |
| 因太空碰撞等意外而导致的返回 | 否 | 是 |
返回式航天器的设计以安全可控地回到地面为目的。由于在目前的技术条件下返回大气层时航天器的速度极高，因此非破坏性返回的过程一般需要有特殊的措施来保护航天器避免受到气动力加热和震动、冲击等损害。由于载人航天一定有航天员返回地面的过程，因此这一过程也成为载人航天中风险较高的环节之一。
歷史
[编辑]這種雙層隔熱板概念在1920年由羅伯特·戈達德提出，他說："流星進入大氣層的速度高達每秒30英里，但內部依然寒冷。因此，假如再返回物的表面覆蓋一層抗高溫（不易變質及難熔解）的物質後再用一層不太會導熱的耐高溫物質，這樣物體表面就不會受到太多的侵蝕"（節錄）。[來源請求]
而第一次實際應用到此系統是在洲际彈道導彈的重返速度增加所導致的气动力加热。早期的彈道導彈，如V2火箭，並沒有此問題。而中程彈道導彈，如蘇聯的R-5（有1200km）的射程，就需要陶瓷複合材料來保護。首個洲際彈道導彈（ICBM，射程達8000至12000km），則已正式進入了現代保護材料的時代。在美國，這技術是由艾姆斯研究中心（NASA Ames）任職的哈立·朱利安·艾倫率先開發。而蘇聯的Yuri A. Dunaev也曾在列寧格勒物理技術研究所開發類似的技術。
飛行器的形狀
[编辑]鈍形飛行器
[编辑]美國國家航空諮詢委員會的哈立·朱利安·艾倫、阿爾弗雷德·J·艾格斯在1951年發現了鈍形（高阻）防热大底具有最佳效果。因為返回式航天器的气动加熱與阻力係數成反比，即阻力愈大，熱負荷愈低。钝形重返舱使得气体不能快速离开，成为气垫层隔开了冲击波与加热振动层，使得大部分热空气不再直接接触重返舱，热能保持在冲击波气体中并在大气层中扩散。
艾倫和艾格斯的發現，最初被視為軍事秘密，但於1958年出版。鈍形理論的設計成為可行的隔熱板，都體現在水星、雙子星和阿波羅太空艙，使宇航員返回火熱的地球大氣層時仍生存。蘇聯的R-7洲際彈道導彈於1957年使用尖鼻的彈頭成功首次試射，但擊中目標區10公里以外，因而改為鈍鼻的彈頭。蘇聯的隔熱層由多層玻璃纖維與石棉textolite組成。
重返飞行器形状
[编辑]球形或球形部分
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1950年代到1960年代，易于从理论上用Fay-Riddell方程建模分析。[3] 当时没有高速计算机，高速空气动力学还处于萌芽阶段。苏联的东方飞船、上升飞船、以及火星、金星探测器的下降舱；美国的阿波罗飞船指令舱都采用了球形防热大底。阿波罗飞船重返时的攻角−27°，升阻比0.368[4]。联盟飞船、月球取样返回探测器、双子座飞船、水星号飞船都是如此设计。即使这少量的升力也使得从弹道式重返8-9g的峰值加速度减小到4-5g，同时大大减少了峰值气动加热。
球-圆锥形
[编辑]
球-圆锥形是指截头圆锥与球形部分的结合。这具有更好的动态稳定性。
美国最早的该构形的重返舱是通用电气于1955年开发的Mk-2 RV。使用了基于金属防热大底的辐射热防护系统(TPS)。Mk-2作为武器投射系统具有很大缺陷，由于低弹道系数使得其在上层大气中飞行时间太长，产生一股金属蒸汽尾流，极易于被雷达发现。
通用电气研制的下一代重返舱是Mk-6，采用尼龙酚醛材质的防热大底，其效果非常好以至于可以大大减小锥体半角到12.5°. Mk-6的重返质量3360 kg，长3.1m。随着核武器小型化与烧蚀材料的进展，重返舱变得更轻、锥体半角减小到10°-11°.
美国日冕侦察卫星是第一种非武器战斗部的重返舱。1959年2月28日首次发射。 February 1959).
不同于军事目的的返回舱，空间探测器的返回舱采取更大的锥体半角显然可以减少对烧蚀材料的需要，降低死重[來源請求]。伽利略号探测器的下降舱的锥体半角达到了45°，海盗号火星着陆舱的锥体半角达到70°
双锥形
[编辑]双锥形是在球形-锥形上又增加了一个截头锥形。这具有非常好的升阻比，达到1.0以上。
非轴对称形状
[编辑]用于载人重返舱，如采用三角翼的航天飞机、暴风雪号航天飞机，以及升力体如X-23 PRIME。
激波层气体物理
[编辑]防热大底设计的经验规则是：激波层气体峰值温度（开尔文为单位）等于进入速度（单位m/s）。例如，宇宙飞船以7.8 km/s进入大气层，激波层气体峰值温度7800 K。因为动能增加与速度的平方成正比，而气体的比热容随温度而大幅度增加（这不同于固体在通常条件下可以假设比热容不变）。
在典型的重返温度，激波层大气是被电离与解离的。这种化学解离必须一些物理模型以描述激波层的热与化学性质。对于设计防热大底的航宇工程师有4种气体基本物理模型：
理想气体模型
[编辑]几乎所有的航太工程师在本科时学过理想气体模型。大部分的理想气体方程与对应的表与图在NACA Report 1135中给出。[5] NACA Report 1135的摘要经常出现在热动力学课本的附录被航宇工程师熟悉。
理想气体理论非常精巧并在设计飞机时非常有用，但它假定气体是化学惰性的。从飞机设计角度，大气温度低于550 K时可以假定气体为惰性的。气温550 K时理想气体理论开始出现问题而气温超过2000 K将不再适用，这时防热大底设计者必须使用真实气体模型。
再入角
[编辑]再入角的大小直接影响到航天器在大气层里的航程。若再入角过小，航天器可能只在大气层边缘掠过而无法进入；若过大则会受到较大的空气阻力，可能会使气动力加热严重进而导致航天器燃烧。一般无人航天器再入大气层时，再入角在3°~8°之间，不会超过10°。[6]为使最大制动过载不超过人体所能耐受的限度（10倍重力加速度），以第一宇宙速度再入的载人飞船必须以小于3°的再入角进入大气层。
参考文献
[编辑]- ^ re-entry - 重返. terms.naer.edu.tw. [2021-12-13]. （原始内容存档于2021-12-13）.
- ^ atmospheric entry - 進入大氣. terms.naer.edu.tw. [2021-12-13]. （原始内容存档于2021-12-13）.
- ^ Fay, J. A.; Riddell, F. R. Theory of Stagnation Point Heat Transfer in Dissociated Air (PDF). Journal of the Aeronautical Sciences. February 1958, 25 (2): 73–85 [2009-06-29]. doi:10.2514/8.7517. （原始内容 (PDF Reprint)存档于2005-01-07）.
- ^ Hillje, Ernest R., "Entry Aerodynamics at Lunar Return Conditions Obtained from the Flight of Apollo 4 (AS-501)," NASA TN D-5399, (1969). (PDF). [2014-11-07]. （原始内容存档 (PDF)于2020-11-11）.
- ^ Equations, tables, and charts for compressible flow. NACA Annual Report (NASA Technical Reports). 1953, 39 (NACA-TR-1135): 611–681 [2014-11-08]. （原始内容存档于2020-09-17）.
- ^ 赵玉晖; 侯锡云; 刘林. 月球返回轨道再入角变化特征. 飞行器测控学报: 75-79 –通过中国知网.
