TW200823108A - Aircraft safety system - Google Patents
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- TW200823108A TW200823108A TW095143198A TW95143198A TW200823108A TW 200823108 A TW200823108 A TW 200823108A TW 095143198 A TW095143198 A TW 095143198A TW 95143198 A TW95143198 A TW 95143198A TW 200823108 A TW200823108 A TW 200823108A
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200823108 九、發明說明：200823108 IX. Description of invention:
肛福^明所屬技糊"領域^ I 發明領域 本發明係有關一設計成可提高意外事件中乘客存活機 5會之飛機安全設施系統。 發明背景 自從1903年12月17曰在凱悌霍克(Kitty Hawk)作出首 次成功的動力駕駛式飛行以來，乘客藉由飛機旅行已成為 10 常態。藉由飛機設計到飛機材料等廣泛領域中之工程師、 科學家及製造廠的顯著進步已經得以製造及生產出能夠攜 載乘客與貨物橫越全球之飛機。全球的乘客數係從1965年 的1億7千7百萬成長到估計2〇〇〇年的33億人。 雖然空中旅行藉由多重備用及安全設施系統一般可說 15是安全且可靠，自從1908年9月17日首次致命墜機以來全球 各地的墜機已經喪失成千上萬的性命。儘管駕駛盡力試圖 找出地面或海上的適當降落位址且在駕駛訓練及飛機精密 性方面作出改良依然如此。 已經提出不同解決方案來致力解決這些飛機相關的喪 2〇生事件。譬如，邱(Chiu)的美國專利案6,382,563號係揭露一 具有一可切斷式外殼及位居殼内的複數個個別乘客艙之飛 機。在緊急事件中，外殼被切斷且個別乘客艙係分離，各 艙配備有獨立的氧氣供源及降落傘。各艙分段隨後能夠在 其各別的降落傘底下漂浮至地面，藉以拯救乘客生命。 5 200823108 類似地，戴門（Diamond)的美國專利案4 699,336號係教 ^一飛機乘客隔室，其含有一用以當飛機有墜機之虞時自 機身射出乘客隔室之機構。然後降落傘能夠使乘客隔室漂 浮至地面。 5 些先前技藝解決方案之一常見問題係在於其不具有 有效的規模。客機不斷增加尺寸，空中巴士(Airbus)最新的 設計A380具有近似80m的翼展，且能夠在兩分離的甲板上 攜載超過550名乘客。具有分段式囊件的飛機之另一相關問 題係在於此系統需要複製的降落傘系統且需要能夠分離各 10囊件之複雜的技術（譬如，火箭及/或雷射），亦需要用於降 落傘系統及分離機構之顯著空間。這當然轉而導致重量及 製造成本的增加。 【發明内容：！ 發明概要 15 為此，本發明提供一飛機，包含： -機身’其具有-上可駕歇機身部分及—下可投棄機 身部分，下可投棄機身部分係在正常飛行期間可拆離式接 合於上可駕驶機身部分； -上翼結構，其與上可駕敬機身部分相聯結； 20 -下翼結構’其與下可投棄機身部分相聯結； 可釋放式接合部件，其用以可拆離式接合下可投棄機 身部分及上可駕駛機身部分； 其中該可釋放式接合部件係可在飛行中致動以使上可 駕駛機身部分及下可投棄機身部分能夠彼此脫離；及 6 200823108 其中可在缺乏下可投棄機身部分之情形下飛行上可駕 駛機身部分。 飛機較佳為一固定式翼客機。上可駕駛機身部分較佳 包括一乘客隔室，且進一步包括一尾段及尾引擎。 5 本發明的一實施例中，尾段係可拆離式接合至上可駕 駛機身部分且可在一緊急狀況中脫離。 下可投棄部分可有利地能夠在從上可駕驶機身部分分 離之後獨立地飛行。下可投棄機身部分較佳進一步包括主 要起落架、燃料箱及貨物儲存隔室。 10肖佳實施例中’下可投棄部分係進-步包括-適可將 下可投棄部分引導至預定安全降落位置之全球定位系統。 此等安全降落位置可有利地預先程式化在全球定位系統中 使得下可投棄機身部分在分離之後受到自動地引導。安全 =洛位置較佳係遠離建築密集區域且可能包括指定的機 15 %、或海面。 上翼結構較佳係包括燃料儲存部件，其用於在自下可 投棄機身部分分離後提供上可駕駛機身部分所使用之燃 20 本發明的-實施例中，上翼結構係在正常飛行期間彼 ^下翼結構中且作為用於飛機之—單經合併翼結構。下 ==可有利地m可釋放式真空部件可釋放式安裝至 口;。構。可釋放式真空部件較佳係與可釋放式接合部件 5夕化以此夠在〆緊急狀況中同步化拆離上及下翼結 構0 7 200823108 本毛明的-替代性實施例中，上翼結構及下翼結構係 在正常飛行期間彼此分離。 上可駕馼機身部分較佳具有輔助控制器及儀器，適可 在下可投棄機身部分已自上可駕駛機身部分分離之後容許 5 一人^駛上可駕駛機身部分。辅雜彻及儀器係為分 在刀離之月;!係為用來駕駛飛機之主控制器及儀器之外 的裝^。上可駕驶機身部分可具有一其中設有輔助控制器 儀。之辅助駕驶區域，其係為分離且身為下可投棄機身 部分分離之前駕駛所佔據以駕駛飛機的主駕駛艙之外的裝 10 置。 可釋放式接合部件可有利地包括一陣列的可釋放式接 合機構。各可釋放式接合機構可包括一可拆離式接合的爪 或夾扣總成。 本發明的-實施例中，可釋放式接合部件係包含一具 15有5員之釋放螺栓，及_夹扣總成。爽扣總成係可包含一 對的爪#可移動於一其中使釋放螺检的頭被接合之經接 a位置及其中使釋放螺栓的頭被釋放之釋放位置之間。 上可駕駛機身部分係可包括辅助起落架，故_旦下可 投棄機身部分被釋放則可降落上可駕駛機身部分。上可 20駛機身部分可進一步包括材料以改善其浮力。 — 本發明的-實施例中，上可駕驶機身部分係包括水上 飛機形狀的船殼幾何結構以在一旦下可投棄機身部分 分離時用以辅助降落在水上。 子 本發明的替代性實施例中，可釋放式接合部件可包含 8 200823108 一Γ軌配置以當可釋放式接合部件被致動時使上可駕駛 機身部分及下可投棄機身部分能夠相對於彼此滑動且變成 脫離。 滑動執配置可有利地包括一齒條及齒輪機構，及至少 5 -對互補的滾子藉以能夠在上可駕馼機身部分訂可投棄 機身部分之間作滑動動作。滑動軌配置較佳進—步包括一 制動機構，其被啟動時係防止上可駕敬機身部分及下可投 棄機身部分相對於彼此滑動，且當解除致動時係能夠使上 可駕駛機身部分及下可投棄機身部分分離。 10圖式簡單說明 現在參照圖式僅藉由示範性範例來插述本發明，其中： 弟1Α至1C圖分別顯示本發明的經修改客機之第一實 施例在分離前之正視圖、側視圖及俯視圖； 第2Α及2Β圖分別顯示處於經分離組態中之第^至^ 15 圖的經修改客機之正視圖及側視圖； 第2C圖顯示處於經分離組態中之飛機的一上可駕驶部 分的俯視平面圖； 第2D圖顯示第2C圖所示的飛機之上可駕駛部分的翼 沿著線2D-2D之橫剖側視圖； 弟2E圖顯不處於經分離組悲中之經修改飛機的下可投 棄部分之俯視平面圖； 第3圖顯示經修改飛機沿著第1B圖的線4A-4A之部分 示意橫剖側視圖； 第4圖顯示第3圖所示之一可釋放式接合機構的詳細橫 9 200823108 剖側視圖； 第5A圖顯示類似於第3圖的部分示意橫剖視側圖圖，其 中上可駕駛部分及下可投棄部分處於經分離組態中； 第5B圖顯示第5A圖的一可釋放式接合機構處於經分 5 離組態中之細節； 第6A至6C圖分別顯示第4圖的可釋放式接合機構處於 經接合組態中之立體圖、側視圖及橫剖正視圖； 第7A至7C圖分別顯示第4圖的可釋放式接合機構處於 釋放位置中之立體圖、側視圖及橫剖正視圖； 10 第8圖顯示一位於飛機的上及下部分之間的密封總成 之詳細剖視立體圖； 第9A圖顯示飛機的一翼沿著第1C圖的線9A-9A之部分 示意橫剖側視圖； 第9B圖顯示第9A圖的翼，其中上及下翼部分處於一經 15 分離組態中； 第9C圖顯示一用以輔助將上及下翼部分固持在一起之 真空歧管； 第10A至10C圖分別顯示本發明的可釋放式接合機構 之第二實施例處於經接合位置中之立體圖、細部圖及部分 20 橫剖側視圖； 第11A及11B圖分別顯示第10A圖的一可釋放式接合機 構處於經脫離位置中之立體圖及部分橫剖側視圖； 第12A至12D圖分別顯示本發明的另一實施例之一低 高度客機在分離之前的等角圖、正視圖、側視圖及橫剖正 10 200823108 視圖， 第13A至13D圖分別顯示第12A至12D的飛機在分離期 間之等角圖、正視圖、側視圖及橫剖正視圖； 第14A至14D圖分別顯示本發明的另一實施例之一高 5 高度客機在分離之前的等角圖、正視圖、側視圖及橫剖正 視圖， 第15A至15D圖分別顯示第14A至14D的飛機在分離期 間之等角圖、正視圖、側視圖及橫剖正視圖； 第16A至16C圖顯示本發明的另一實施例之一安全設 10 施滑翔機在分離之前的等角圖、正視圖、側視圖及橫剖正 視圖； 第17A至17D圖分別顯示第16A至16C的飛機在分離期 間之等角圖、正視圖、及側視圖； 第18A至18C圖分別顯示本發明的另一實施例之一雙 15 引擎飛機在分離之前的等角圖、正視圖、及側視圖； 第19A至19C圖分別顯示第18A至18C的飛機在分離期 間之等角圖、正視圖、及側視圖； 第20A至20C圖分別顯示本發明的另一實施例之一經 嵌套引擎高高度飛機在分離之前的等角圖、正視圖、及側 20 視圖； 第21A至21C圖分別顯示第20A至20C的飛機在分離期 間之等角圖、正視圖、及側視圖； 第22A至22D圖分別顯示本發明的另一實施例之一高 高度發射器（launcher)在分離之前的等角圖、正視圖、側視 11 200823108 圖及橫剖正視圖； 第23A至23D圖分別顯示第22A至22D的發射器在分離 期間之等角圖、正視圖、及側視圖； 第23E圖顯示第22A至22D圖的ARAVS在飛行至太空中 5 期間之橫剖侧視圖， 第24A至24D圖分別顯示本發明的另一實施例之一超 音波三角翼飛機在分離之前的等角圖、正視圖、側視圖及 橫剖正視圖； 第25A至25D圖分別顯示第24A至24D的飛機在分離期 10 間之等角圖、正視圖、側視圖及橫剖正視圖； 第26A至26C圖分別顯示本發明的另一實施例之一穿 音速三角翼飛機在分離之前的等角圖、正視圖、及側視圖； 第27A至27C圖分別顯示第26A至26C的飛機在分離期 間之等角圖、正視圖、及側視圖； 15 第28A至28C圖分別顯示本發明的另一實施例之一三 階段飛機在分離之前的等角圖、正視圖、及側視圖； 第29A1至29C圖分別顯示第28A至28C的飛機在分離期 間之等角圖、正視圖、及側視圖； 第29D至29G圖分別顯示無機尾引擎模組之第29A1至 2〇 29C圖的飛機之上部分的等角圖、正視圖、及側視圖； 第30A至30C圖分別顯示本發明的另一實施例之一雙 翼飛機在分離之前的等角圖、正視圖、及側視圖； 第31A至31C圖分別顯示第30A至30C的雙翼飛機在分 離期間之等角圖、正視圖、及側視圖； 12 200823108 第32A1至32D圖分別顯示本發明的另一實施例之一雙 翼高高度發射器在分離之前的等角圖、正視圖、側視圖、 及橫剖正視圖； 第33A至33E圖分別顯示第32A1至32D的發射器在分離 5 期間之等角圖、正視圖、側視圖、及橫剖正視圖； 第34A至34C圖分別顯示本發明的另一實施例之一曲 柄-V(crank-V)飛機在分離之前的等角圖、正視圖、及側視 圖， 第35A至35C圖分別顯示第34A至34C的飛機在分離期 10 間之等角圖、正視圖、及側視圖； 第36A1至36C圖分別顯示本發明的另一實施例之一箱 翼(box-wing)飛機在分離之前的等角圖、正視圖、及側視圖； 第37A至37C圖分別顯示第36A1至36C的飛機在分離期 間之等角圖、正視圖、及側視圖； 15 第38A至38D圖分別顯示本發明的另一實施例之一併 入有一滑動釋放機構之高高度客機在分離之前的等角圖、 正視圖、側視圖及橫剖正視圖； 第38E及38F圖顯示第38A至38D的欄杆系統在分離之 前的放大橫剖正視圖； 20 第39A至39C圖分別顯示第38A至38D的飛機在分離期 間之等角圖、側視圖及橫剖正視圖； 第40A至40C圖分別顯示第38A至38D的飛機在分離之 後的專角圖、側視圖、及側視圖， 第40D及40E圖顯示第38A至38D的欄杆系統在分離之 13 200823108 後的放大橫剖正視圖； 第41圖顯示第38A至38D的飛機在分離之前的放大橫 剖正視圖而其中制動系統被致動； 第42A至42B圖顯示第38A至38D的飛機之經致動制動 5系統在飛行期間的放大橫剖正視圖； 第43圖顯示一用以分離飛機的上及下部分之液壓釋放 機構。 較佳實施例之詳細說明 1〇 首先參照第1Α至1C圖，顯示將在正常操作環境下飛行 之根據本發明的一實施例之一經修改客機1〇。客機1〇具有 一機身12、一組的翼14、翼安裝式引擎16、18及一容置有 一尾引擎21之尾段2〇。 經修改客機1〇係沿著一可密封介面22被分成一上可駕 u駛部分說及_下可投棄部分_。如下文更詳細地描述， 發生-緊急狀況時，下可投棄部分10B構形為可自上可駕驶 部分10A脫離。 上可駕駛部分10A係包括一乘客隔室24及出口門28。飛 機的-主駕駛艙26係設置於上可駕駛部分1〇A的前部且含 20有當上可駕馱部分10A及下可投棄部分ι〇Β在正常操作環 境中附接域此時絲駕駛飛機之控㈣及儀器。 翼I4由上翼部分MA及下翼部分MB構成。上翼部分 W被攜載於飛機的上可駕駛部分上，而下翼部分⑽ 連同翼引擎16、18被攜載於飛機的下可投棄部分應上。在 14 200823108 正常飛行狀況下，上及下翼部分i4A及14B係以第9A圖所示 的方式穩固地彼此接合以界定機翼14。第2E圖顯示下翼部 分14B如何設有用以容納上翼部分14A之互補形狀的凹部 40。下可投棄部分中亦設有前及後貨物隔室30，及用以容 5置經修改客機1〇的習知起落架（未圖示）之前及主起落架隔 室32 、 34 。 現在參照第2A及2B圖，第1A至ic圖之經修改客機以 的上可駕駛部分及下可投棄部分l〇A、10B係顯示處於經脫 離或經分離組態中。一般係由於諸如飛機1〇的下可投棄部 10分10B之一組件發生損害等一緊急狀況而發生脫離。飛機1〇 的下可投棄部分10B係包括攜載於下翼部分14B中之前及 後貨物隔室30、主翼燃料箱及翼引擎16、μ。當貨物隔室 30完全地裝載且主翼燃料箱充滿時，下可投棄部分Mg可代 表飛機10上之整體負載質量的幾乎一半。因此，在一緊急 15狀況中投棄飛機10A的下部分1〇B將有利於藉此提高上可 駕驶部分安全地降落之機會。 上可駕駛部分10A係包括上翼部分14A、尾段2〇、尾引 擎2卜及被攜載於上翼部分14A中之翼安裝式燃料箱37。第 2D圖顯示設置於上翼部分14A中之翼安裝式燃料箱37的橫 20剖視圖。上可駕馱部分1〇A係能夠以駕驶、組員及乘客的整 個全體呈《立地飛行。上可駕駛部分說亦設有輔助前及 主起落架35、36藉以容許可駕駛部分1〇A作出緊急降落。 上可駕駛部分1GA較佳亦設有編控制ϋ及儀器，諸如 轉向輛、高度計、無線電及引擎控制器等，以在自飛機1〇 15 200823108 的下可：棄部分分離後容許—人駕驶上可駕驶部分 。輔助控㈣較佳係與上及下部分、應分離之前 用來駕駛經修改客卿之主控㈣及絲呈賴立。輔助 控制器及儀祕可有利地設置於—㈣與主駕駛麟分離 且獨立之伽駕驶M27巾。辅叫駛區域加意地顯示 於第1C圖中成為—設置於主駕敬艙％後方之小房間。因 此’如果主駕聽26受損或駕駛員例如由於煙霧累積而無 法接達時’上可駕駛部分H)A可自下可投棄部分應分離且 可自輔助駕駛27飛行。 〇 s代性實施例中’上可駕驶部分10A可在自下可投棄部 分10B分離之後利用主控制器及儀器或辅助控制器及儀器 予以駕駛。在此例中，輔助控制器可設置於主駕駛艙26中 或可如上述設置於一輔助駕駛區域27中。或者，並未提供 輔助控制器及儀器且上可駕駛部分1〇Α在自下可投棄部分 15 10B分離之後利用主駕駛艙26中的主控制器及儀器予以駕 驶0 現在參照第3圖，以橫剖面顯示飛機上之機身的上及下 部分10A及10B。上可駕駛部分10A係包括被附裝至飛機之 上可駕駛部分10A的一下地板44之乘客座位42。一充氣室46 20 被界定於下地板44及飛機的下可投棄部分10B上之一上天 花板48之間。充氣室46較佳容納設置於各別殼體52中之6個 可釋放式接合機構50。可釋放式接合機構較佳係配置成位 居經修改客機1〇的機首中之一前對、與機翼14相鄰之一中 央對、及位居飛機10的機尾中之一後對。 16 200823108 可釋放式接合機構50係具有在脫離之前於大部份狀況 下將飛機10的上可駕駛及下可投棄部分10A及10B固持在 一起之作用。各殼體52係穩固地安裝在飛機的上部分i〇A 内，如第3圖所示。各殼體52的一下凸緣54係穩固地抵靠住 5 飛機1〇的下部分10B之上天花板48。各可釋放式接合機構50 係包括一對的液壓致動器56而其作用在一爪總成58上以將 一釋放螺栓62的經擴張頭60固持在一經接合位置中。釋放 螺栓62穩固地安裝至下可投棄部分10B的上天花板48。當令 爪總成58在液壓致動器56影響下開啟時，釋放螺栓62的頭 10 60係被釋放造成飛機的整體可投棄部分ιοΒ自上可駕駛部 分10A釋放並掉落離開上可駕駛部分1〇八。 第5A圖更詳細地顯示可釋放式接合機構5〇已被致動之 後上可駕驶部分10A自下可投棄部分ιοΒ分離。第5B圖係為 一可釋放式接合機構50處於經脫離位置中之放大細部圖。 15釋放螺栓62已自可釋放式接合機構50之爪總成58被釋放並 已經自可釋放式接合機構5〇分離。 第6A至6C及7A至7C圖中更詳細地顯示可釋放式接合 機構50之操作。首先參照第6八至6(：圖，可釋放式接合機構 50係包括一由在各端接合之第一及第二分開的三角形板 20 64A及64B所形成之框架料。一中央銷66係橋接板的頂點， 而一對内及外爪68及70係可旋轉地安裝至中央銷66。第6A 至6C圖顯示處於經接合位置中之可釋放式接合機構，爪係 界定了拘限插座72而在其内保持困住釋放螺栓的經擴張頭 60。爪68及70之相對側係可_地安裝至各別槓桿臂顺 17 200823108 76。槓桿们4及76_可_地絲至在㈣糊往復移 動之液壓鎚78的端點以形成液壓致動器56。各紅體80的基 底係安裳至框架64的_部分。外槓桿臂料係具有位於框 架的相對基端處之外固定式樞軸點86，及可供内横桿臂74 5及76以及致動益78配置成繞其樞轉之内樞軸點88。馬蹄鉤 總成90係可樞轉地安裳至横桿臂科的上部分，而線纔或結 合桿92自各馬蹄鉤延伸。 現在參照第7A至7C圖，顯示位於一釋放位置中之可釋 放式接合機構50。釋放位置中，液壓致動器％被致動以使 10枢軸點88往上位移，藉以造成内及外槓桿臂74及84 、及76 及84朝向彼此樞轉’藉以驅迫内及外爪砧及川開啟且使釋 放螺栓62的經擴張頭60自由。所有6個可釋放式接合機構的 同時釋放係將造成飛機1〇的下可投棄部分1〇]3自上可駕駛 部分10A被投棄。在液壓故障之事件中，經由在自馬蹄鉤9〇 15延伸的線纜或桿92上往上拉來實行人工致動。可經由位居 飛機的上可駕駛部分10A之下地板44上方的安穩位置中之 對應的鬆緊螺旋扣總成來達成此作用。 將瞭解人工及液壓致動機構皆受到具有有限強制控制 特性之不同保全系統，以確保不會發生意外或未經授權的 20 釋放。 第8圖顯示一包括一密封墊片94之密封總成用以提供 飛機的上及下部分l〇A、10B之間的一可密封及緩衝介面。 墊片94係自一適當彈性材料形成，且如同飛機的下部分1〇B 之上天花板48及飛機的上部分i〇A之一最下方地板96具有 18 200823108 適當輪廓，以准許兩部分相對於彼此之受圍堵式運動。一 層膜材料設置於最下方地板96的上方，其在第3圖中示意地 顯示為98，以在一水面降落事件中增加飛機的上可駕駛部 分之浮力。 5 現在參照第9A及9B圖，更詳細地顯示翼μ的上及下翼 部分14A及14B。下翼部分14B係配合於設有諸如清除邊蓋 102或其他清除機構等燃料清除器之燃料箱ι〇〇，以容許一 旦自上翼部分14A分離即自下翼部分14B清除燃料。這使得 下可投棄部分10B在衝擊時爆炸的機會達到最小。上翼部分 1〇 14A貼切地嵌套在下翼部分14B内。提供一更清楚地顯示於 第9C圖中之真空歧管配置1〇4以確保上及下翼部分在正常 飛行狀況期間被固持為彼此接合。真空歧管丨〇2係包括一系 列之交替的肋108及與一真空歧管11〇導通的通路1〇6。肋可 為被界定於下翼部分14B的上表面或上翼部分14A的下表 15 面中之褶縐形式。 真空歧管110可容許更易生成一真空。真空歧管11〇係 進給至通路106中以使真空歧管110中的壓力影響通路1〇6 中的壓力。利用多重通路由於一通路中的壓力失效將不會 影響其他通路中的壓力因而改良了上翼部分14A附接至下 2〇 翼部分UB之可靠度。 可釋放式接合機構50致動時，真空被同時地釋放以容 許上及下翼部分14A、14B以第9B圖所示的方式彼此脫離^ 將瞭解上及下翼部分14A、14B兩者皆配合於邊蓋及副翼總 成。特定言之，上翼部分HA係具有相對較充分的辅助邊蓋 19 200823108 及副翼以在自下可投棄部分10B分離之後可容許飛行及控 制飛機之上可駕駛部分10A。 現在參照第1()A及腦圖，示—可釋放式接合機構 110的第二實施例用以實行釋放螺检62的—受控制脫離。第 5 10A至10C圖顯示位於經接合位置中之可釋放式接合機構 10可釋放式接合機構11〇係包括—對的液壓致動器⑴及 114 ’其臂係可樞轉地連接至自旋轉掣子丨觀伸之各別横 桿臂U6及118。掣子12〇係攜裁有在一插座箱126内所界定 的對應圓柱形腔穴丨糊移動之拱形止動部122。拱形止動 1〇部122可移動於第1〇C圖所示其中使其留置經釋放螺栓62之 經擴張頭60的-經接合位置、與第UB圖所示其中使致動紅 體112及114_㈣以將止動部⑵姉朝向彼此使釋放 螺栓62自由之一經釋放位置之間。 第11A及11B圖顯示處於釋放&置中之可釋放式接合 15機構110止動部120已經旋轉故使拱形止動部122移至一釋 放位置且釋放螺栓62驗擴張頭60自插座箱126釋放。 本發明的另一實施例係顯示於第12a_12d及13A_13d 圖中。首先參照第12A-12D圖，顯示如上述之一嵌套式翼飛 機1〇’包含上可駕駛部分10A及下可投棄部分1〇B。如第12d 20及13D圖清楚地顯示，上可駕駛部分1〇A係併入一包括一v 型船殼200之水上飛機船殼幾何結構。v型船殼2〇〇係提供一 水切割邊緣，且連帶具有位於船殼2〇〇兩側上之喷灑條片 202,降低上可駕駛部分說的水衝擊負荷並辅助成功的降 低及水上的漂浮。上可駕駛部分1〇A亦設有一供海面上漂浮 20 200823108 及穩定所用之浮力段202。浮力段202較佳包含抗水、抗火、 且抗衝擊藉以符合安全要求之低密度複合泡緯。 第12A-12D及13A-13D圖的飛機較佳進一步包括身為 下可投棄部分10B的一部分之一對的可佈署水平安定面及 5升降舵204。這些水平安定面及升降舵係供縱向俯仰穩定度 所用且使用於下可投棄部分10B自上可駕駛部分1〇A分離 之後。 現在參照第14A-14D及15A-15D圖，顯示飛機1〇的另一 實施例。此實施例中，飛機10適可作高高度飛行及分離。 10飛機10具有欲套式翼14，及如上述之上及下部分1〇八、 10B。此實施例中，上可駕駛部分10A及下可投棄部分1〇B 係形成有經過設計可在正常操作期間及緊急飛行期間維持 内部壓力之橢圓形橫剖面。上可駕駛部分10A因此能夠在高 高度自下可投棄部分分離同時兩者部分保持被加壓。飛機 15 10係進一步設有一整流罩2〇6及支撐件208形式之振動消減 特性。 第16A-16C及17A-17D圖係顯示本發明的另一實施 例，其中飛機10的上可駕駛部分10A經過修改可在一緊急狀 況的事件中以一安全滑翔機飛行。此實施例中，自下可投 20棄部分10B分離之後，上可駕駛部分10A保持一駕駛控制式 載具、具有輕重量、且只供滑翔用途。上可駕駛部分10A 係包括一輔助動力單元及包括可佈署腹側水平安定面及升 降舵204等緊急降落所需要的相關設備。第17D圖中顯示上 可駕馱部分10A處於其滑翔組態中。飛機1〇的大部份重量係 21 200823108 ;匕括下翼部分14B、主引擎16、18、尾段2G及尾引擎21 之可投棄部分10B，如第17C圖清楚地顯示。 第UA-ISC及19A-19C圖顯示與上文對於第16A_16C及 17A 17D圖所描述者相類似之另一滑翔實施例。此實施例 5中，飛機10適可作高高度飛行及分離，且併入有通常屬於 小至中航裎客機之雙尾引擎21A、21B。 第2〇A_20C及21A-21C圖顯示本發明的另一實施例。此 實^例中’尾引擎21係在上可駕駛部分1()A中被欲套於飛機 的機尾機身12中。肷套式尾引擎21係包括相聯結的空氣攝 1〇入部210。此實施例的飛機10係適可作高高度飛行及分離且 包括一併入一對方向舵之獨特的水平穩定器212。水平穩定 态212設置於飛機10的下可投棄部分10B上。在一緊急的事 件中，尾段20連同水平穩定器214係與飛機的下部分10B被 投棄，如第21C圖清楚地顯示。 15 現在參照第22A_22D及23A-23E圖，顯示應用至一低地 軌道再進入載具發射器220之本發明的一實施例。此實施例 係併入較早實施例之嵌套式翼組態14A、14B。 此實施例中，較早實施例的上可駕駛部分1〇A係採行一 自主火箭輔助式載具段(ARAVS) 220A的形式，而下可投棄 20部分10B採行一可重覆使用式高高度運送發射器220B的形 式。ARAVS 220A係為一預定將可能身為太空設備及/或乘 客的一酬載221攜載至大空中之自主太空載具。 兩段220A係以單一飛機被發射且利用液體（或可再裝 填固體）燃料222在發射的第一階段中將攸升至一預定發射 22 200823108 同度在此點’造成可重覆使用式發射器220B自ARAVS 220A刀離且回到其原始目的地。可重覆使用式發射FIELD OF THE INVENTION The present invention relates to an aircraft safety system designed to enhance passenger survival in an accident. BACKGROUND OF THE INVENTION Since the first successful power-driving flight at Kitty Hawk on December 17th, 1903, passengers have become the norm by air travel. Significant advances in engineers, scientists and manufacturing plants in a wide range of fields, from aircraft design to aircraft materials, have enabled the manufacture and production of aircraft capable of carrying passengers and cargo across the globe. The number of passengers worldwide has grown from 177 million in 1965 to 3.3 billion in the estimated 2 years. Although air travel is generally safe and reliable with multiple standby and safety facilities, the crashes around the world have lost thousands of lives since the first fatal crash on September 17, 1908. Despite driving to try to find the appropriate landing site on the ground or at sea and to improve on driving training and aircraft precision. Different solutions have been proposed to address these aircraft-related mourning events. For example, U.S. Patent No. 6,382,563 to Chiu discloses a aircraft having a severable outer casing and a plurality of individual passenger compartments located within the casing. In an emergency, the enclosure is cut and individual passenger compartments are separated, each compartment being equipped with an independent oxygen supply and a parachute. Each compartment segment can then float to the ground under its respective parachute to save the lives of passengers. 5 200823108 Similarly, U.S. Patent No. 4,699,336 to the United States teaches an aircraft passenger compartment that includes a mechanism for ejecting a passenger compartment from the fuselage when the aircraft has a crash. The parachute then floats the passenger compartment to the ground. One of the common problems with some of the prior art solutions is that they do not have an effective scale. Airliners continue to increase in size, and Airbus's latest design, the A380, has a wingspan of approximately 80m and is capable of carrying more than 550 passengers on two separate decks. Another related problem with aircraft with segmented capsules is that the system requires a duplicated parachute system and requires complex techniques (eg, rockets and/or lasers) capable of separating the 10 capsules, as well as for parachutes. Significant space for the system and the separation mechanism. This of course leads to an increase in weight and manufacturing costs. [Summary of the Invention:! SUMMARY OF THE INVENTION To this end, the present invention provides an aircraft comprising: - a fuselage' having an upper body portion and a lower body portion, the lower fuselage portion being detachable during normal flight Disengaged to the upper driveable fuselage section; - upper wing structure that is coupled to the upper steerable fuselage section; 20 - lower wing structure 'which is coupled to the lower disposable fuselage section; releasable engagement component The detachable joint can be used to dispose of the fuselage portion and the upper drivable fuselage portion; wherein the releasable joint member is actuatable in flight to enable the upper driveable fuselage portion and the lower disposable machine The body parts can be separated from each other; and 6 200823108 which can fly the body part in the absence of the fuselage part. The aircraft is preferably a fixed wing passenger aircraft. The upper drivable fuselage portion preferably includes a passenger compartment and further includes a tail section and a tail engine. In an embodiment of the invention, the tail section is detachably coupled to the upper driveable fuselage section and is detachable in an emergency. The lower disposable portion can advantageously fly independently after being separated from the upper drivable fuselage portion. The lower disposable fuselage portion preferably further includes a main landing gear, a fuel tank, and a cargo storage compartment. In the 10th embodiment, the lower disposable portion includes a global positioning system that is adapted to direct the disposable portion to a predetermined safe landing position. Such safe landing positions can advantageously be pre-programmed in the global positioning system such that the lower disposable fuselage sections are automatically guided after separation. Safety = The Lok position is preferably far from the densely populated area and may include 15% of the designated machine, or the sea surface. The upper wing structure preferably includes a fuel storage component for providing combustion of the upper drivable fuselage portion after separation from the lower fuselable fuselage portion. In the embodiment of the invention, the upper wing structure is normal. During the flight, in the lower wing structure and as a single-merged wing structure for the aircraft. Lower == advantageously the releasable vacuum component can be releasably mounted to the port; Structure. The releasable vacuum member is preferably coupled to the releasable engagement member 5 to synchronize the detachment of the upper and lower wing structures in an emergency. 0 7 200823108 In an alternative embodiment, the upper wing The structure and the lower wing structure are separated from each other during normal flight. The upper controllable body portion preferably has an auxiliary controller and an instrument, and the lower part of the fuselage portion can be separated from the upper body of the driveable body. The auxiliary and the instrument are divided into the month of the knife; the system is used to drive the aircraft's main controller and equipment. The upper drivable fuselage section may have an auxiliary controller therein. The auxiliary driving area is separate and can be thrown away from the main cockpit of the aircraft that is occupied by the driver before the separation. The releasable engagement member can advantageously include an array of releasable engagement mechanisms. Each releasable engagement mechanism can include a detachably engageable jaw or clip assembly. In the embodiment of the invention, the releasable engaging member comprises a release bolt having 15 members and 5 members, and a clip assembly. The cool buckle assembly can include a pair of jaws # movable between a position in which the head that releases the thread check is engaged, and a release position in which the head that releases the bolt is released. The upper driveable fuselage section may include an auxiliary landing gear, so that the disposable fuselage section may be released to land on the driveable fuselage section. The upper body portion of the fuselage may further include materials to improve its buoyancy. - In the embodiment of the invention, the upper drivable fuselage section comprises a hull geometry of a watercraft shape to assist in landing on the water once the fuselable portion is disengaged. In an alternative embodiment of the invention, the releasable engagement member can include 8 200823108 a rail configuration to enable the upper and lower disposable fuselage portions to be opposite when the releasable engagement member is actuated Sliding on each other and becoming detached. The sliding configuration may advantageously include a rack and pinion mechanism, and at least 5 - pairs of complementary rollers for enabling a sliding action between the upper steerable body portion and the disposable body portion. Preferably, the sliding rail arrangement includes a braking mechanism that is activated to prevent the upper steerable fuselage portion and the lower disposable fuselage portion from sliding relative to each other, and is capable of driving upward when deactivated The fuselage part and the lower part can be separated from the fuselage part. BRIEF DESCRIPTION OF THE DRAWINGS The present invention will now be described by way of example only, with reference to the drawings, wherein: FIG. 1 to FIG. 1C respectively show a front view and a side view of a first embodiment of the modified passenger aircraft of the present invention before separation. And the top view; the second and second figures respectively show the front view and the side view of the modified passenger aircraft in the separated configuration of the ^ to ^ 15; Figure 2C shows the one of the aircraft in the separated configuration A top plan view of the driving portion; Figure 2D shows a cross-sectional side view of the wing of the drivable portion above the aircraft shown in Figure 2C along line 2D-2D; the 2E picture is not modified in the sadness of the separated group A top plan view of the lower disposable portion of the aircraft; Figure 3 shows a cross-sectional side view of the modified aircraft along line 4A-4A of Figure 1B; Figure 4 shows one of the releasable joints shown in Figure 3 Detailed view of the mechanism 9 200823108 cross-sectional side view; Figure 5A shows a partial schematic cross-sectional side view similar to Figure 3, wherein the upper drivable part and the lower disposable part are in the separated configuration; Figure 5B shows a releasable joint of Figure 5A The details are in the configuration of the split 5; the 6A to 6C are respectively a perspective view, a side view and a cross-sectional front view of the releasable joint mechanism of Fig. 4 in the engaged configuration; Figs. 7A to 7C A perspective view, a side view, and a cross-sectional elevation view, respectively, showing the releasable engagement mechanism of Figure 4 in the release position; 10 Figure 8 shows a detailed cross-sectional perspective view of the seal assembly between the upper and lower portions of the aircraft Figure 9A shows a cross-sectional side view of a portion of the aircraft along line 9A-9A of Figure 1C; Figure 9B shows the wing of Figure 9A with the upper and lower wing portions in a 15 separate configuration; Figure 9C shows a vacuum manifold for assisting in holding the upper and lower wing portions together; Figures 10A through 10C show perspective views of the second embodiment of the releasable engagement mechanism of the present invention in the engaged position, FIG. 11A and FIG. 11B are respectively a perspective view and a partial cross-sectional side view showing a releasable engagement mechanism of FIG. 10A in a disengaged position; FIGS. 12A to 12D respectively show the present invention Another An isometric view, a front view, a side view, and a cross-section of a low-height passenger aircraft in an embodiment prior to separation. 200823108 views, Figures 13A to 13D show isometric views of the aircraft of the 12A to 12D during separation, Front view, side view and cross-sectional front view; Figures 14A to 14D respectively show an isometric view, a front view, a side view and a cross-sectional front view of the high 5 height passenger aircraft of the other embodiment of the present invention before separation. 15A to 15D are respectively an isometric view, a front view, a side view, and a cross-sectional front view of the aircraft of the 14A to 14D during the separation; FIGS. 16A to 16C show a safety device 10 of another embodiment of the present invention. An isometric view, a front view, a side view and a cross-sectional front view of the glider before separation; Figures 17A to 17D respectively show an isometric view, a front view, and a side view of the aircraft of the 16A to 16C during separation; Figures 18A to 18C show an isometric view, a front view, and a side view, respectively, of a dual 15-engine aircraft of another embodiment of the present invention before separation; Figures 19A to 19C show the aircraft of the 18A to 18C during separation, respectively. Isometric view And FIGS. 20A to 20C respectively show an isometric view, a front view, and a side 20 view of another embodiment of the present invention before the separation by the nested engine high-height aircraft; FIGS. 21A to 21C respectively show An isometric view, a front view, and a side view of the aircraft of FIGS. 20A to 20C during separation; FIGS. 22A to 22D respectively show equiangular heights of a high-altitude launcher of another embodiment of the present invention before separation Figure, front view, side view 11 200823108 Figure and cross-sectional front view; Figures 23A to 23D show the isometric, front and side views of the emitters of 22A to 22D, respectively, during separation; Figure 23E shows Cross-sectional side view of the ARAVS of Figures 22A to 22D during flight to space 5, Figures 24A to 24D respectively showing an isometric view, front view of the ultrasonic delta wing aircraft of another embodiment of the present invention prior to separation , side view and cross-sectional front view; Figures 25A to 25D show the isometric, front, side and cross-sectional elevation views of the aircraft of the 24A to 24D in the separation period, respectively; Figures 26A to 26C show Another embodiment of the present invention An isometric view, a front view, and a side view of a Sonic Delta aircraft before separation; Figures 27A through 27C show isometric, front, and side views of the aircraft of Figures 26A through 26C, respectively; Figures 28A to 28C respectively show an isometric view, a front view, and a side view of a three-stage aircraft of another embodiment of the present invention before separation; Figures 29A1 to 29C show the aircraft of the 28A to 28C during separation, respectively. Isometric view, front view, and side view; Figures 29D through 29G show isometric, front, and side views, respectively, of the upper portion of the aircraft of Figures 29A1 through 2〇29C of the inorganic tail engine module; 30A to 30C respectively show an isometric view, a front view, and a side view of the biplane of another embodiment of the present invention before separation; FIGS. 31A to 31C show the separation of the biplanes of the 30A to 30C, respectively. Isometric, front view, and side view of the period; 12 200823108 Figures 32A1 to 32D respectively show an isometric view, a front view, a side view of a bi-wing high-altitude launcher of another embodiment of the present invention prior to separation And cross-sectional front view; 33 A to 33E respectively show an isometric view, a front view, a side view, and a cross-sectional front view of the emitters of the 32A1 to 32D during the separation 5; FIGS. 34A to 34C respectively show one of the other embodiments of the present invention The isosceles, front view, and side view of the crank-V aircraft before separation. Figures 35A to 35C show the isometric, front view, and the 10, respectively, of the aircraft in the 34A to 34C separation period. And side views; Figures 36A1 to 36C respectively show an isometric view, a front view, and a side view of a box-wing aircraft of another embodiment of the present invention before separation; Figures 37A to 37C show respectively An isometric view, a front view, and a side view of the aircraft of Figures 36A1 to 36C during separation; 15 Figures 38A to 38D respectively show a high altitude passenger aircraft incorporating a sliding release mechanism in accordance with another embodiment of the present invention. Previous isometric, front, side and cross-sectional elevation views; Figures 38E and 38F show enlarged cross-sectional elevation views of the railing system of sections 38A to 38D prior to separation; 20 Figures 39A to 39C show section 38A, respectively Isometric view, side view of the aircraft to 38D during separation Cross-sectional front view; Figures 40A to 40C show the angle diagram, side view, and side view of the aircraft after the separation of 38A to 38D, respectively. Figures 40D and 40E show the railing system of the 38A to 38D in separation 13 A magnified cross-sectional elevation view after 200823108; Figure 41 shows an enlarged cross-sectional elevation view of the aircraft of 38A to 38D prior to separation with the brake system actuated; Figures 42A to 42B show the aircraft of 38A to 38D An enlarged cross-sectional elevation view of the actuating brake 5 system during flight; Figure 43 shows a hydraulic release mechanism for separating the upper and lower portions of the aircraft. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 1A Referring first to Figures 1 to 1C, there is shown a modified guest aircraft 1 according to an embodiment of the present invention that will fly in a normal operating environment. The passenger aircraft 1 has a fuselage 12, a set of wings 14, a wing-mounted engine 16, 18, and a tail section 2 that houses a tail engine 21. The modified passenger aircraft 1 is divided along a sealable interface 22 into an upper driveable portion and a lower disposable portion. As described in more detail below, in the event of an emergency, the lower disposable portion 10B is configured to be detachable from the upper drivable portion 10A. The upper drivable portion 10A includes a passenger compartment 24 and an exit door 28. The main cockpit 26 of the aircraft is disposed at the front of the upper drivable portion 1A and includes 20 the upper steerable portion 10A and the lower disposable portion ι attached to the field in a normal operating environment. Control of the aircraft (four) and instruments. The wing I4 is composed of an upper wing portion MA and a lower wing portion MB. The upper wing portion W is carried on the upper drivable portion of the aircraft, while the lower wing portion (10) is carried along with the wing engines 16, 18 on the lower disposable portion of the aircraft. Under normal flight conditions, 14 200823108, the upper and lower wing portions i4A and 14B are securely engaged with each other to define the wing 14 in the manner shown in Figure 9A. Figure 2E shows how the lower wing portion 14B is provided with a recess 40 for receiving the complementary shape of the upper wing portion 14A. The front and rear cargo compartments 30 are also provided in the lower disposable portion, and the conventional landing gears (not shown) and the main landing gear compartments 32, 34 are used to accommodate the modified passenger aircraft. Referring now to Figures 2A and 2B, the upper and lower disposable portions of the modified passenger aircraft of Figures 1A through ic are shown in a disengaged or separated configuration. Generally, the detachment occurs due to an emergency such as damage to a component of the lower disposable portion 10 minutes 10B of the aircraft. The lower disposable portion 10B of the aircraft includes the front and rear cargo compartments 30, the main wing fuel tank and the wing engines 16, μ carried in the lower wing portion 14B. When the cargo compartment 30 is fully loaded and the main wing fuel tank is full, the lower disposable portion Mg can represent almost half of the overall load mass on the aircraft 10. Therefore, the abandonment of the lower portion 1B of the aircraft 10A in an emergency 15 condition will facilitate the opportunity to safely land the upper drivable portion. The upper drivable portion 10A includes an upper wing portion 14A, a tail portion 2A, a tail engine 2, and a wing-mounted fuel tank 37 carried in the upper wing portion 14A. Fig. 2D shows a cross-sectional view of the wing 20 of the wing-mounted fuel tank 37 disposed in the upper wing portion 14A. The upper part of the system can be used for driving, group members and passengers. The upper drivable part said that there are also auxiliary front and main landing gears 35, 36 to allow the drivable part 1A to make an emergency landing. Preferably, the upper driving part 1GA is also provided with a control device and an instrument, such as a steering gear, an altimeter, a radio and an engine controller, etc., to be available from the aircraft 1〇15 200823108: the abandoned part is separated and allowed to be driven by the person. The driving part. Auxiliary control (4) The preferred system and the upper and lower parts should be used to drive the modified main control (4) and the silk display. The auxiliary controller and the instrument secret can advantageously be set to - (4) separate from the main driving lining and independently drive the M27 towel. The auxiliary driving area is intentionally displayed in the small room in the 1st figure, which is set to the rear of the main driving cabin. Therefore, if the main driver 26 is damaged or the driver cannot access it due to, for example, smoke accumulation, the upper drivable portion H)A can be separated from the lower disposable portion and can be self-assisted driving 27 to fly. In the s generation embodiment, the upper drivable portion 10A can be driven by the main controller and the instrument or the auxiliary controller and instrument after the separation from the lower disposable portion 10B. In this example, the auxiliary controller may be disposed in the main cockpit 26 or may be disposed in an auxiliary driving area 27 as described above. Alternatively, the auxiliary controller and the instrument are not provided and the upper drivable portion 1 is driven by the main controller and instrument in the main cockpit 26 after the separation from the lower disposable portion 15 10B. Referring now to FIG. 3, The cross section shows the upper and lower portions 10A and 10B of the fuselage on the aircraft. The upper drivable portion 10A includes a passenger seat 42 attached to the lower floor 44 of the upper drivable portion 10A of the aircraft. An plenum 46 20 is defined between the lower floor 44 and one of the ceiling panels 48 on the lower disposable portion 10B of the aircraft. The plenum 46 preferably houses six releasable engagement mechanisms 50 disposed in the respective housings 52. The releasable engagement mechanism is preferably configured to be in front of one of the heads of the modified passenger aircraft, one center pair adjacent to the wing 14, and one of the tails of the aircraft 10. . 16 200823108 The releasable engagement mechanism 50 has the function of holding the upper drivable and lower disposable portions 10A and 10B of the aircraft 10 together under most conditions prior to disengagement. Each housing 52 is securely mounted within the upper portion i〇A of the aircraft, as shown in FIG. The lower flange 54 of each of the housings 52 securely abuts against the ceiling 48 above the lower portion 10B of the aircraft. Each releasable engagement mechanism 50 includes a pair of hydraulic actuators 56 that act on a jaw assembly 58 to retain the expanded head 60 of a release bolt 62 in an engaged position. The release bolt 62 is securely mounted to the upper ceiling 48 of the lower disposable portion 10B. When the pawl assembly 58 is opened under the influence of the hydraulic actuator 56, the head 10 60 of the release bolt 62 is released causing the overall disposable portion of the aircraft to be released from the upper drivable portion 10A and dropped off the upper drivable portion 1 Eight. Figure 5A shows in more detail that the releasable engagement mechanism 5 has been actuated and the upper drivable portion 10A is separated from the lower disposable portion ιο. Figure 5B is an enlarged detail view of a releasable engagement mechanism 50 in a disengaged position. The release bolt 62 has been released from the jaw assembly 58 of the releasable engagement mechanism 50 and has been detached from the releasable engagement mechanism 5〇. The operation of the releasable engagement mechanism 50 is shown in more detail in Figures 6A through 6C and 7A through 7C. Referring first to Figures 6-8 to 6 (the figure, the releasable engagement mechanism 50 includes a frame material formed by first and second divided triangular plates 20 64A and 64B joined at each end. A central pin 66 The apex of the bridge plate, and a pair of inner and outer jaws 68 and 70 are rotatably mounted to the center pin 66. Figures 6A through 6C show the releasable engagement mechanism in the engaged position, the jaws defining the restraint socket 72. The expanded head 60 is held therein to trap the release bolt. The opposite sides of the claws 68 and 70 can be mounted to the respective lever arms cis 17 200823108 76. The levers 4 and 76_ can be grounded to (d) The end of the hydraulic hammer 78 that reciprocates to form a hydraulic actuator 56. The base of each red body 80 is attached to the portion of the frame 64. The outer lever arm has a position other than the opposite base end of the frame. The fixed pivot point 86, and the inner crossbar arms 74 5 and 76 and the actuation benefit 78 are configured to pivot about the inner pivot point 88. The horseshoe hook assembly 90 is pivotally mounted to the cross The upper part of the arm arm, and the line or the coupling rod 92 extends from each horseshoe hook. Referring now to Figures 7A to 7C, the display is located at one The releasable engagement mechanism 50 in the release position. In the release position, the hydraulic actuator % is actuated to displace the 10 pivot point 88 upwardly, thereby causing the inner and outer lever arms 74 and 84, and 76 and 84 to face each other Pivoting 'by virtue of urging the inner and outer anvils and the opening of the shaft and freeing the expanded head 60 of the release bolt 62. The simultaneous release of all six releasable engagement mechanisms will result in a lower disposable portion of the aircraft 1〇] 3 The upper driveable portion 10A is discarded. In the event of a hydraulic failure, manual actuation is performed by pulling up on a cable or rod 92 extending from the horseshoe hook 9〇15. The corresponding elastic screw-sleeve assembly in the stable position above the floor 44 below the driving portion 10A achieves this effect. It will be appreciated that both the manual and hydraulic actuation mechanisms are subjected to different security systems with limited mandatory control characteristics to ensure that no occurrence will occur. Unexpected or unauthorized 20 release. Figure 8 shows a seal assembly including a gasket 94 for providing a sealable and cushioning interface between the upper and lower portions 10A, 10B of the aircraft. 94 series from one When the elastic material is formed, and the ceiling 48, like the lower portion 1B of the aircraft, and the lowermost floor 96 of the upper portion i〇A of the aircraft, have an appropriate contour of 18 200823108 to permit the two parts to be trapped relative to each other. Movement. A layer of film material is disposed above the lowermost floor 96, which is shown schematically at 98 in Figure 3 to increase the buoyancy of the upper drivable portion of the aircraft during a surface landing event. 5 Reference is now made to Figures 9A and 9B. Figure, showing the upper and lower wing portions 14A and 14B of the wing μ in more detail. The lower wing portion 14B is fitted to a fuel tank ι provided with a fuel remover such as a cleaning side cover 102 or other cleaning mechanism to allow once The fuel is removed from the lower wing portion 14B when the upper wing portion 14A is separated. This minimizes the chance that the disposable portion 10B will explode upon impact. The upper wing portion 1A 14A is nested in the lower wing portion 14B. A vacuum manifold arrangement 1〇4, shown more clearly in Figure 9C, is provided to ensure that the upper and lower wing portions are held in engagement with one another during normal flight conditions. The vacuum manifold 丨〇 2 includes a series of alternating ribs 108 and a passage 1 〇 6 that is in communication with a vacuum manifold 11 . The ribs may be in the form of pleats defined in the lower surface of the lower wing portion 14B or in the lower surface of the upper wing portion 14A. The vacuum manifold 110 allows for a more vacuum to be created. The vacuum manifold 11 is tethered into the passage 106 such that the pressure in the vacuum manifold 110 affects the pressure in the passage 1〇6. The use of multiple passages will not affect the pressure in the other passages due to pressure failure in one passage and thus improve the reliability of attachment of the upper wing portion 14A to the lower 2 flap portions UB. When the releasable engagement mechanism 50 is actuated, the vacuum is simultaneously released to allow the upper and lower wing portions 14A, 14B to disengage from each other in the manner shown in Figure 9B. ^ It will be understood that both the upper and lower wing portions 14A, 14B are mated. In the side cover and aileron assembly. In particular, the upper wing portion HA has a relatively sufficient auxiliary side cover 19 200823108 and ailerons to allow for flight and control of the drivable portion 10A above the aircraft after separation from the lower disposable portion 10B. Referring now to the first () A and brain diagram, a second embodiment of the releasable engagement mechanism 110 is used to effect the controlled release of the release screw 62. Figures 5A through 10C show the releasable engagement mechanism 10 in the engaged position. The releasable engagement mechanism 11 includes a pair of hydraulic actuators (1) and 114' whose arms are pivotally coupled to the self-rotating掣 丨 丨 之 之 之 之 横 横 横 横 横 横 横 横 横 横The tweezers 12 carry an arched stop 122 that moves in a corresponding cylindrical cavity as defined in a socket 126. The arched stop 1 can be moved to the engaged position of the expanded head 60 through which the release bolt 62 is retained as shown in FIG. 1C, and the red body 112 is actuated as shown in FIG. And 114_(d) to position the stoppers (2) toward each other such that one of the release bolts 62 is free between the released positions. 11A and 11B show the releasable engagement 15 mechanism 110 in the release & centering stop 120 has been rotated so that the arched stop 122 is moved to a release position and the release bolt 62 is inspected from the socket box 126 release. Another embodiment of the invention is shown in Figures 12a-12d and 13A_13d. Referring first to Figures 12A-12D, it is shown that one of the nested wing aircraft 1'' comprises an upper drivable portion 10A and a lower disposable portion 1B. As clearly shown in Figures 12d 20 and 13D, the upper drivable portion 1A is incorporated into a seaplane hull geometry including a v-shaped hull 200. The v-shaped hull 2 provides a water-cut edge with associated spray strips 202 on either side of the hull 2 to reduce the water impact load on the upper drivable portion and assist in successful reduction and water Floating. The upper drivable part 1A also has a buoyancy section 202 for floating on the sea surface 20 200823108 and for stabilization. The buoyancy section 202 preferably comprises a low density composite weft that is resistant to water, fire, and impact to meet safety requirements. The aircraft of Figures 12A-12D and 13A-13D preferably further includes an deployable horizontal stabilizer and a fifth elevator 204 that are part of a portion of the lower disposable portion 10B. These horizontal stabilizers and elevators are used for longitudinal pitch stability and are used after the lower disposable portion 10B is separated from the upper drivable portion 1A. Referring now to Figures 14A-14D and 15A-15D, another embodiment of an aircraft 1 显示 is shown. In this embodiment, the aircraft 10 is adapted for high altitude flight and separation. The aircraft 10 has a wraparound wing 14, and upper and lower portions 1-8, 10B as described above. In this embodiment, the upper drivable portion 10A and the lower disposable portion 1B are formed with an elliptical cross section designed to maintain internal pressure during normal operation and during emergency flight. The upper drivable portion 10A is thus capable of separating portions at a high height from the lower disposable portion while the portions remain pressurized. The aircraft 15 10 is further provided with a vibration reduction characteristic in the form of a fairing 2〇6 and a support member 208. Figures 16A-16C and 17A-17D show another embodiment of the present invention in which the upper drivable portion 10A of the aircraft 10 is modified to fly in a safe glider in the event of an emergency. In this embodiment, after the separation of the lower disposable portion 10B, the upper drivable portion 10A maintains a driving control type carrier, has a light weight, and is only used for gliding purposes. The upper drivable portion 10A includes an auxiliary power unit and associated equipment including an emergency landing such as a deployable ventral horizontal stabilizer and a lift rudder 204. The upper controllable portion 10A is shown in Fig. 17D in its gliding configuration. The majority of the weight of the aircraft is 21 200823108; the lower wing portion 14B, the main engine 16, 18, the tail section 2G and the disposable portion 10B of the tail engine 21 are clearly shown as shown in Fig. 17C. Figures UA-ISC and 19A-19C show another gliding embodiment similar to that described above for Figures 16A-16C and 17A 17D. In this embodiment 5, the aircraft 10 is adapted for high altitude flight and separation, and incorporates two-tailed engines 21A, 21B which are typically small to medium-sized passenger aircraft. Figures 2A-20C and 21A-21C show another embodiment of the present invention. In this embodiment, the tail engine 21 is intended to be fitted in the tail body 12 of the aircraft in the upper drivable portion 1 () A. The clamshell tail engine 21 includes an associated air intake unit 210. The aircraft 10 of this embodiment is suitable for high altitude flight and separation and includes a unique horizontal stabilizer 212 incorporating a pair of rudders. The horizontal steady state 212 is disposed on the lower disposable portion 10B of the aircraft 10. In an emergency event, the tail section 20 along with the horizontal stabilizer 214 is discarded from the lower portion 10B of the aircraft, as shown clearly in Figure 21C. 15 Referring now to Figures 22A-22D and 23A-23E, an embodiment of the present invention applied to a low earth track re-entry vehicle launcher 220 is shown. This embodiment incorporates the nested wing configuration 14A, 14B of the earlier embodiment. In this embodiment, the upper drivable portion 1A of the earlier embodiment takes the form of an autonomous rocket-assisted vehicle segment (ARAVS) 220A, and the lower portion 20b 10B adopts a reusable type. The high height transports the form of the transmitter 220B. The ARAVS 220A is an autonomous space vehicle that is intended to carry a payload 221 that may be a space device and/or passenger to the air. The two sections 220A are launched in a single aircraft and utilize liquid (or refillable solids) fuel 222 to ramp up to a predetermined launch during the first phase of launch 22 200823108. At the same point 'causes reusable launches The switch 220B is detached from the ARAVS 220A and returns to its original destination. Reusable launch
器 220B 係〇括可佈署腹側水平安定面及升降航加4。分離之後， ARAVS 22GA係利用固體或液體火箭燃料似的—供應源繼 5續其任務進入最高達到近似臟m的低高度執道The 220B system includes a deployable ventral horizontal stabilizer and a hoistway plus 4. After separation, the ARAVS 22GA utilizes a solid or liquid rocket fuel-like supply source to continue its mission to a low altitude of up to approximately dirty m.
。ARAVS 220A係①相於再進人狀況以供可能的乘客飛行之最大安 全性且對於方向性控制利用推力向量控制。 本毛明的另一貫施例係顯示於第24A-24D及25A-25D 圖中。此貫施例係為一併入一超音波三角翼設計以供一般 1〇大於1馬赫數速度之超音波三角翼飛機。此實施例中，在機 尾列板226中採用雙垂直穩定器。機尾列板226係與飛機10 的下部分1GB之翼14B-體地形成。上可駕駛部分1〇A係包 括分離之前被嵌套於下部分14B的翼14B中之水平安定面 及升降1204。可駕駛部分i〇A係包括如對於較早實施例所 15描述分離後之自主飛行所需要的特性並可進一步包括一用 於動力滑翔之引擎（未圖示）。 第26A-26C及27A-27C圖顯示本發明的另一實施例。本 發明的此實施例除了此飛機為穿音速三角翼之外概括係與 上文對於第24A-24D及25A-25D圖所描述者相同。翼μ對於 2〇位於0·8至1.2馬赫數範圍中之穿音速予以最佳化。再者，可 駕馱部分10A可視需要選用地包括一用於動作滑翔之引擎 (未圖示），且或者具有分離之後自主飛行所需要的特性如同 對於較早實施例所描述。 本發明的另一實施例顯示於第28A-28C及29A1-29G圖 23 200823108 中。此實施例係為上述實施例之進一步延伸，並包括第三 可佈署飛機模組。特別參照第28C、29C、及29E圖，此實 施例的飛機10係包含三個模組—如前述之上可駕歇部分 10A及下可投棄部分1(^，及一用以容置尾引擎21之額外的 5可佈署機尾引擎模組10C。尾引擎21具有兩功能，第一係為 在正常飛行期間提供推力，第二係為使用於動力飛行的_ 緊急狀況之案例中藉以到達一目的地及安全的降低。在尾 引擎或方向舵故障之事件中，機尾模組1〇c可自可駕駛部分 10A被投棄。在此例中，上可駕駛部分1〇A將如較早所述以 10 一高高度翼滑翔機進行。 上可駕駛部分1 〇 A係併入一水上飛機船殼幾何結構，如 對於第12D圖所描述，藉以使其適合藉由一尾引擎或不具有 尾引擎而在水上作出成功的降落或漂浮。並且，上可駕駛 部分係包括未被嵌套於下可投棄部分1〇B的翼ΐ4β中之一 15對的翼14A。藉由將高翼14A設置於上可駕駛部分上係可幫 助上部分l〇A安全地降落於水上。. The ARAVS 220A is one-phase in re-entry for maximum safety for possible passenger flight and thrust vector control for directional control. Another example of the present invention is shown in Figures 24A-24D and 25A-25D. This embodiment is an ultrasonic delta wing aircraft incorporating an ultrasonic delta wing design for a general speed greater than 1 Mach. In this embodiment, a dual vertical stabilizer is employed in the tail column 226. The tailgate 226 is integrally formed with the lower 1GB wing 14B of the aircraft 10. The upper drivable portion 1A includes a horizontal stabilizer and a lift 1204 that are nested in the wing 14B of the lower portion 14B before separation. The drivable portion i 〇 A includes the characteristics required for the autonomous flight after separation as described in the earlier embodiment 15 and may further include an engine for power gliding (not shown). Figures 26A-26C and 27A-27C show another embodiment of the present invention. This embodiment of the invention is the same as that described above for Figures 24A-24D and 25A-25D except that the aircraft is a transonic delta wing. The wing μ is optimized for the sound transmission speed of 2 〇 in the range of M·8 to 1.2 Mach. Further, the controllable portion 10A optionally includes an engine for action gliding (not shown), and may have the characteristics required for autonomous flight after separation as described for earlier embodiments. Another embodiment of the invention is shown in Figures 28A-28C and 29A1-29G Figure 23 200823108. This embodiment is a further extension of the above embodiment and includes a third deployable aircraft module. With particular reference to Figures 28C, 29C, and 29E, the aircraft 10 of this embodiment includes three modules - such as the above-described rideable portion 10A and the lower disposable portion 1 (^, and one for receiving the tail engine An additional 5 deployable tail engine module 10C. The tail engine 21 has two functions, the first system is to provide thrust during normal flight, and the second system is used in the case of emergency use for power flight. A destination and a safe reduction. In the event of a tail engine or rudder failure, the tail module 1〇c can be discarded from the drivable portion 10A. In this example, the upper drivable portion 1A will be as early as. The upper glider is carried out with a high-altitude wing glider. The upper drivable part 1 〇A is incorporated into a seaplane hull geometry, as described for Figure 12D, thereby making it suitable for use with a tail engine or without tail The engine makes a successful landing or floating on the water. And, the upper drivable part includes 15 pairs of wings 14A that are not nested in the lower wing portion 4B of the lower portion 1B. By setting the high wing 14A On the upper part of the driver can help the upper part l〇A Full landing on the water.
弟川 A-30C 及 31A-31C 、 〜^ wI他例，亦 20 即:適可用純高度、未加壓分離之雙翼飛機。雙翼飛機 1〇係包括-上可駕驶部分隐及下可投棄部分_。上可駕 驶部分1GA係包括—對的非絲式高翼UA。高翼14Α被三 為在上可駕駛部分降糾水上日村提供 二 到最小。上可駕驶部㈣係包括二= 行«安全乘客運送所描述之全部所需要特性， 匕括一水上飛機船殼幾何結構。 24 200823108 本發明的另一實施例係顯示於第32A1_32D及 33A2_33E圖中。這些圖式顯示與第22A-22D及23A-23E圖所 描述者相似之一低地執道發射器22〇。此實施例不同處在於 翼14A及14B並未嵌套而是以一雙翼組態設置。翼14A、14B 5的此配置係提供發射器220中之經改良的結構性效率。 第34A-34C及35A-35C圖顯示一曲柄-V飛機10作為本 發明的另一實施例。曲柄-V飛機1〇係包括上可駕駛部分i〇A 及下可投棄部分10B。上可駕駛部分i〇A係設有沿著下部分 10B的翼14B附接於中途之一對的翼14A。翼14A以翼鎖固機 10構228可拆離地緊固至下翼14B，如第35A-35C圖清楚地顯 示。翼鎖固機構228係與可釋放式接合機構5〇同步化以使分 離程序期間的安全性達到最大。曲柄_v組態因為所有表面 皆用於揚升故改良了飛機的效能，且結構性效率受到改 善。曲柄-V飛機亦併入兩個經加壓橢圓形缸體及振動消減 15 特性如第14D圖所示以供高高度操作。 第36A1-36C及37A-37C圖顯示本發明的另一實施例， 亦即一適可用於低高度、非加壓分離之箱翼飛機。箱翼飛 機係與第30A-30C及31A-31C圖的雙翼飛機相同但添加了 被併入上可駕駛部分10A的翼14A與可投棄部分10B的翼 20 14B之間的左及右支柱230。藉由此配置，飛機受到結構上 及空氣動力學上的改良。負荷及勁度(stiffness)係由支柱230 轉移於翼14A、14B之間，其藉由盡量降低翼梢處的旋渦效 應而空氣動力性地降低了所引發的拖髮力。 本發明的此實施例進一步包括被併入飛機的支柱230 25 200823108 中之水面降落穩定器232。水面降落穩定器232採行一安裝 在各支柱230上之可充氣袋的形式而其在分離程序之後且 恰在一緊急降落之前被充氣。水面降落穩定器232連同上可 駕駛部分10A之水上飛機船殼幾何結構係改良一緊急水面 5降落期間上部分10A的效能。支柱230亦可併入一可藉以不 再需要垂直穩定器及來自機尾機身的前導邊緣之方向舵 (未圖示）。 上述的實施例中，已經經由快速釋放接合機構5〇且在 散套翼飛機案例中包括一額外的真空系統來達成飛機的上 10可駕駛部分10A及下可投棄部分10B之接合。 本發明的另一實施例顯示於第38A至38F、39A至39C、 40A至40E、4卜42A至42B圖中。此實施例中，接合機構5〇 係被一具有相聯結的制動機構27〇之滑動軌系統250所取 代0 15 滑動軌系統250由一齒條252及齒輪254配置所組成，其 具有一致動器256，及位於飛機的一側之一對滾子，包含上 溝槽狀滾子258及下滾子260，以及飛機相對側上之上滾子 262及下滾子264。滾子258、260、262、264係設置於飛機 的下可投棄部分10B之壁266上，如第38E及38F圖清楚地顯 2〇 示。壁266係延伸經過設置於上可駕駛部分i〇A的基底中之 開孔267。上可駕駛部分10A及下可投棄部分10B分離之 前，各別對的滾子258、260以及262、264係設置為與上可 駕駛部分10A的一下水平壁268相鄰且位居上方及下方。一 密封件269設置於上可駕駛部分10A及下可投棄部分10B之 26 200823108 間。 如下文所述由於一旦制動機構270被解除時則上可駕 驶部分10A及下可投棄部分·一般能夠自由地滑動，並非 在所有環境中皆需要齒條252及齒輪254之致動器256。部分 5環境中，需要額外力量來啟動分離程序。額外力量可由致 動器256提供且可包括機械釋放裝置及/或氣動件。亦可藉 由將一真空系統逆轉成正壓力來達成分離。 現在參照如第42A及42B圖最清楚地顯示之制動機構 270，該機構係包括位於飛機各側之一對的制動系統，各制 10動系統包括一制動墊、屐、及振動消減材料。參照圖式， 顯示上溝槽狀制動墊272及下制動墊274，及位於飛機的相 對側處之上制動墊276及下制動墊278。 制動機構270進一步包括制動致動器280、緊固件282、 及壓縮彈簧284。當制動機構270被啟動時，壓縮彈簧284作 15用係為固持住緊固件282使得下制動墊274、278壓抵住且接 合上可駕駛部分10A的壁268。當制動器由致動器28〇解除 時’彈簧282被釋放以容許釋放緊固件282及制動墊。 氣機的正$ 4呆作期間，制動機構270隨時皆被啟動藉以 在上可駕駛部分10A與下可投棄部分10B之間維持一穩固 20接合。需要時，譬如在一緊急狀況中，制動機構被解除以 容許上可駕駛部分10A及下可投棄部分1〇B相對於彼此滑 動且變成分離。在具有嵌套式翼之飛機中，施加至翼14A、 14B之真空系統係與制動系統同時地被釋放。 圖中，顯示下可投棄部分10B往上可駕駛部分i〇A前方 27 200823108 滑動但請瞭解下可投棄部分亦可往上可駕驶部分i〇a後方 滑動。 第43圖顯示一液壓釋放機構520,其可用來強迫上機身 口P刀與下機身部分之間的_初始分離。液壓致動器5⑻及框 5轉#506可安裝在上機身部分1〇A上，而凸耳5〇8則位居下機 身部分10B上。當兩部分、應接合，樞轉臂鄉之一釣 形端510係扣持住凸耳5〇8。當分離被引發時，液壓致_ 500在方向502中拉動樞轉臂5〇6。臂5〇6係沿上機身部分上 之點504旋轉，而在拱形方向512中移動。釣形端51〇因此釋 10放凸耳508且在方向514中推押凸耳514,因此推押上及下機 身部分分開。-旦空氣流動於上及下部分之間，兩部分被 強迫進一步分開。一使部分分離之替代性方法係逆轉一真 空壓力以產生一正壓力。 將瞭解此說明書中所揭露及界定之本發明係延伸至文 15字或圖式所提及或得知之個別特性的兩或更多者之所有M 代性組合。所有這些組合係構成本發明之不同替代性態樣。 【圖式簡單明】 第1A至1C圖分別顯示本發明的經修改客機之第一實 施例在分離前之正視圖、側視圖及俯視圖； 20 第2A及2B圖分別顯示處於經分離組態中之第^至^ 圖的經修改客機之正視圖及側視圖； 第2C圖顯示處於經分離組態中之飛機的一上可駕驶部 分的俯視平面圖； 第2D圖顯示第2C圖所示的飛機之上可駕駛部分的翼 28 200823108 沿著線2D-2D之橫剖側視圖； 第2E圖顯示處於經分離組態中之經修改飛機的下可投 棄部分之俯視平面圖； 第3圖顯示經修改飛機沿著第1B圖的線4A-4A之部分 5 示意橫剖侧視圖； 第4圖顯示第3圖所示之一可釋放式接合機構的詳細橫 剖側視圖； 弟5 A圖顯示類似於弟3圖的部分示意橫剖視側圖圖，其 中上可駕駛部分及下可投棄部分處於經分離組態中； 10 第5B圖顯示第5A圖的一可釋放式接合機構處於經分 離組態中之細節； 第6A至6C圖分別顯示第4圖的可釋放式接合機構處於 經接合組態中之立體圖、側視圖及橫剖正視圖； 第7A至7C圖分別顯示第4圖的可釋放式接合機構處於 15釋放位置中之立體圖、側視圖及橫剖正視圖； 第8圖顯示一位於飛機的上及下部分之間的密封總成 之詳細剖視立體圖； 第9A圖顯示飛機的一翼沿著第lc圖的線9A_9A之部分 示意橫剖側視圖； 2〇 第9B圖顯示第9A圖的翼，其中上及下翼部分處於一經 分離組態中； 第9C圖顯示-用以輔助將上及下翼部分固持在一起之 真空歧管； 第10A至10C圖分別顯示本發明的可釋放式接合機構 29 200823108 - 之第二實施例處於經接合位置中之立體圖、細部圖及部分 . 橫剖側視圖； 第11A及11B圖分別顯示第10A圖的一可釋放式接合機 構處於經脫離位置中之立體圖及部分橫剖側視圖； 5 第12A至12D圖分別顯示本發明的另一實施例之一低 高度客機在分離之前的等角圖、正視圖、側視圖及橫剖正 視圖， 第13A至13D圖分別顯示第12A至12D的飛機在分離期 間之等角圖、正視圖、側視圖及橫剖正視圖； 10 第14A至14D圖分別顯示本發明的另一實施例之一高 高度客機在分離之前的等角圖、正視圖、側視圖及橫剖正 視圖， 第15A至15D圖分別顯示第14A至14D的飛機在分離期 間之等角圖、正視圖、側視圖及橫剖正視圖； 15 第16A至16C圖顯示本發明的另一實施例之一安全設 施滑翔機在分離之前的等角圖、正視圖、側視圖及橫剖正 視圖； 第17A至17D圖分別顯示第16A至16C的飛機在分離期 間之等角圖、正視圖、及側視圖； 20 第18A至18C圖分別顯示本發明的另一實施例之一雙 引擎飛機在分離之前的等角圖、正視圖、及側視圖； 第19A至19C圖分別顯示第18A至18C的飛機在分離期 間之等角圖、正視圖、及側視圖； 第20A至20C圖分別顯示本發明的另一實施例之一經 30 200823108 嵌套引擎高高度飛機在分離之前的等角圖、正視圖、及側 視圖， 第21A至21C圖分別顯示第20A至20C的飛機在分離期 間之等角圖、正視圖、及側視圖； 5 第22A至22D圖分別顯示本發明的另一實施例之一高 高度發射器（launcher)在分離之前的等角圖、正視圖、側視 圖及橫剖正視圖； 第23A至23D圖分別顯示第22A至22D的發射器在分離 期間之等角圖、正視圖、及側視圖； 10 第23E圖顯示第22A至22D圖的ARAVS在飛行至太空中 期間之橫剖側視圖； 第24A至24D圖分別顯示本發明的另一實施例之一超 音波三角翼飛機在分離之前的等負圖、正視圖、側視圖及 橫剖正視圖； 15 第25A至25D圖分別顯示第24A至24D的飛機在分離期 間之等角圖、正視圖、側視圖及橫剖正視圖； 第26A至26C圖分別顯示本發明的另一實施例之一穿 音速三角翼飛機在分離之前的等角圖、正視圖、及側視圖； 第27A至27C圖分別顯示第26A至26C的飛機在分離期 20 間之等角圖、正視圖、及側視圖； 第28A至28C圖分別顯示本發明的另一實施例之一三 階段飛機在分離之前的等角圖、正視圖、及側視圖； 第29A1至29C圖分別顯示第28A至28C的飛機在分離期 間之等角圖、正視圖、及側視圖； 31 200823108 第29D至29G圖分別顯示無機尾引擎模組之第29A1至 29C圖的飛機之上部分的等角圖、正視圖、及側視圖； 第30A至30C圖分別顯示本發明的另一實施例之一雙 翼飛機在分離之前的等角圖、正視圖、及側視圖； 5 第31A至31C圖分別顯示第30A至30C的雙翼飛機在分 離期間之等角圖、正視圖、及側視圖； 弟32A1至32D圖分別顯示本發明的另一實施例之一雙 翼高高度發射器在分離之前的等角圖、正視圖、側視圖、 及橫剖正視圖； 10 第33A至33E圖分別顯示第32A1至32D的發射器在分離 期間之等角圖、正視圖、側視圖、及橫剖正視圖； 第34A至34C圖分別顯示本發明的另一實施例之一曲 柄-V(crank-V)飛機在分離之前的等角圖、正視圖、及側視 圖； 15 第35A至35C圖分別顯示第34A至34C的飛機在分離期 間之等角圖、正視圖、及側視圖； 第36A1至36C圖分別顯示本發明的另一實施例之一箱 翼(box-wing)飛機在分離之前的等角圖、正視圖、及側視圖； 第37八至37c圖分別顯示第36ak36C的飛機在分離期 20間之等角圖、正視圖、及側視圖； 第38A至38D圖分別顯示本發明的另一實施例之一併 入有-滑動釋放機構之高高度客機在分離之前的等角圖、 正視圖、側視圖及橫剖正視圖； 第38E及38F圖顯示第38A至38D的攔杆系統在分離之 32 200823108 前的放大橫剖正視圖； 第39A至39C圖分別顯示第38A至38D的飛機在分離期 間之等角圖、側視圖及橫剖正視圖； 第40A至40C圖分別顯示第38A至38D的飛機在分離之 5 後的等角圖、側視圖、及側視圖； 第40D及40E圖顯示第38A至38D的攔杆系統在分離之 後的放大橫剖正視圖； 第41圖顯示第38A至38D的飛機在分離之前的放大橫 剖正視圖而其中制動系統被致動； 10 第42A至42B圖顯示第38A至38D的飛機之經致動制動 系統在飛行期間的放大橫剖正視圖； 第43圖顯示一用以分離飛機的上及下部分之液壓釋放 機構。 【主要元件符號説明】 10…經修改客機 22…可密封介面 10A···上可駕駛部分 24…乘客隔室 10B···下可投棄部分 26…主駕駛艙 10C···機尾模組 27…輔助駕駛區域 12…機身 28…出口門 14A…上翼部分 3〇…前及後貨物隔室 14B…下翼部分 32,34···前及主起落架隔室 16，18···翼安裝式引擎 35,36···輔助前及主起落架 20…尾段 37…翼安裝式燃料箱 21…尾引擎 40···凹部 33 200823108 42…乘客座位 44···下地板 46…充氣室 48…上天花板 50…可釋放式接合機構 52…殼體 54·· ·下凸緣 56，112，114,500〜液壓致動器 58…爪總成 60…經擴張頭 62…釋放螺栓 64…框架 64A，64B…第一及第二分開的 三角形板 66…中央銷 68,70···内及外爪 72…拘限插座 74,76，116，118〜槓桿臂 78…液壓鎚 80…缸體 84…外槓桿臂 88…内極轴點 90…馬蹄鉤總成 92…線纜或結合桿 94…密封墊片 96…最下方地板 98…膜材料 100…燃料箱 102···真空歧管 110…可釋放式接合機構 (fig.l0A，10B)，真空歧管 (fig.9A，9B) 102···清除邊蓋 104···真空歧管配置 106…通路 108···肋 120···旋轉掣子 122…拱形止動部 124···圓柱形腔穴 126…插座箱 200···ν型船殼 202…喷灑條片 204···可佈署腹側水平安定面 及升降舵 206…整流罩 208…支撐件 210···空氣攝入部 212…水平穩定器 34 200823108 220…低地軌道再進入載具發射器 266···壁 220···Α自主火箭輔助式載具 267…開孔 段(ARAVS) 268…下水平壁 220Β···可重覆使用式高高度運 269…密封件 送發射器 270…制動機構 221…酬載 272…上溝槽狀制動墊 222…液體(或可再裝填固體)燃料 274,278…下制動墊 226···機尾列板 276…上制動塾 228···翼鎖固機構 280…制動致動器 230…支柱 282…緊固件 232···水面降落穩定器 284…壓縮彈簧 250···滑動執系統 502…方向 252…齒條 504···點 254…齒輪 506…樞轉臂 256···致動器 508,514…凸耳 258···上溝槽狀滾子 510…釣形端 260,264···下滾子 512…拱形方向 262…上滾子 520…液壓釋放機構 35The brothers A-30C and 31A-31C, ~^wI, and 20 are: suitable for pure-wing, unpressurized biplanes. Biplanes 1 包括 includes - the upper part of the drivable part can be discarded. The upper drive section 1GA includes a pair of non-wire high wing UAs. The high wing 14Α is provided by the third to the minimum in the upper part of the driving section. The upper drivable (4) system includes all the required features described in the two paragraphs «Safe Passenger Transport, including a seaplane hull geometry. 24 200823108 Another embodiment of the present invention is shown in Figures 32A1_32D and 33A2_33E. These figures show a low-level emissive transmitter 22A similar to that described in Figures 22A-22D and 23A-23E. The difference in this embodiment is that the wings 14A and 14B are not nested but are arranged in a two-wing configuration. This configuration of the wings 14A, 14B 5 provides improved structural efficiency in the transmitter 220. Figures 34A-34C and 35A-35C show a crank-V aircraft 10 as another embodiment of the present invention. The crank-V aircraft 1 system includes an upper drivable portion i〇A and a lower disposable portion 10B. The upper drivable portion iA is provided with a wing 14A attached to one of the pair of wings 14B along the lower portion 10B. The wing 14A is detachably fastened to the lower wing 14B by a wing locker 10, as shown clearly in Figures 35A-35C. The wing locking mechanism 228 is synchronized with the releasable engagement mechanism 5 to maximize safety during the separation procedure. The crank _v configuration improves the performance of the aircraft because all surfaces are used for lifting, and the structural efficiency is improved. The Crank-V aircraft also incorporates two pressurized elliptical cylinders and vibration reduction 15 characteristics as shown in Figure 14D for high altitude operation. Figures 36A1-36C and 37A-37C show another embodiment of the present invention, i.e., a box-wing aircraft suitable for low altitude, non-pressurized separation. The wing-wing aircraft is identical to the biplanes of Figures 30A-30C and 31A-31C but adds left and right struts 230 between the wings 14A that are incorporated into the upper drivable portion 10A and the wings 20 14B of the disposable portion 10B. . With this configuration, the aircraft is structurally and aerodynamically improved. The load and stiffness are transferred between the wings 14A, 14B by the struts 230, which aerodynamically reduces the induced drag force by minimizing the vortex effect at the wing tips. This embodiment of the invention further includes a surface landing stabilizer 232 that is incorporated into the struts 230 25 200823108 of the aircraft. The surface landing stabilizer 232 takes the form of an inflatable bag mounted on each of the struts 230 which is inflated after the separation procedure and just prior to an emergency landing. The surface landing stabilizer 232, together with the seaplane hull geometry of the upper drivable portion 10A, improves the performance of the upper portion 10A during an emergency water surface 5 landing. The struts 230 can also incorporate a rudder (not shown) that can no longer require a vertical stabilizer and a leading edge from the tail fuselage. In the above-described embodiment, the engagement of the upper 10 drivable portion 10A and the lower disposable portion 10B of the aircraft has been achieved via the quick release engagement mechanism 5 and including an additional vacuum system in the case of the loose wing aircraft. Another embodiment of the present invention is shown in Figures 38A through 38F, 39A through 39C, 40A through 40E, and 4b 42A through 42B. In this embodiment, the engagement mechanism 5 is replaced by a slide rail system 250 having an associated brake mechanism 27. The slide rail system 250 is comprised of a rack 252 and gear 254 configuration having an actuator. 256, and one of the pair of rollers on one side of the aircraft, includes an upper grooved roller 258 and a lower roller 260, and an upper roller 262 and a lower roller 264 on opposite sides of the aircraft. The rollers 258, 260, 262, 264 are disposed on the wall 266 of the lower disposable portion 10B of the aircraft, as clearly shown in Figures 38E and 38F. The wall 266 extends through an aperture 267 provided in the base of the upper drivable portion i〇A. Before the upper drivable portion 10A and the lower disposable portion 10B are separated, the respective pairs of rollers 258, 260 and 262, 264 are disposed adjacent to the lower horizontal wall 268 of the upper drivable portion 10A and are positioned above and below. A seal 269 is disposed between the upper drivable portion 10A and the lower disposable portion 10B 26 200823108. As described below, the upper steerable portion 10A and the lower disposable portion are generally free to slide once the brake mechanism 270 is released, and the rack 252 and the actuator 256 of the gear 254 are not required in all environments. In Part 5, additional power is required to initiate the separation process. Additional force may be provided by actuator 256 and may include mechanical release means and/or pneumatics. Separation can also be achieved by reversing a vacuum system to a positive pressure. Referring now to the brake mechanism 270, which is most clearly shown in Figures 42A and 42B, the mechanism includes a brake system located on one of the sides of the aircraft, each of which includes a brake pad, cymbal, and vibration abatement material. Referring to the drawings, upper grooved brake pads 272 and lower brake pads 274 are shown, as well as brake pads 276 and lower brake pads 278 at opposite sides of the aircraft. Brake mechanism 270 further includes a brake actuator 280, a fastener 282, and a compression spring 284. When the brake mechanism 270 is activated, the compression spring 284 is used to hold the fastener 282 such that the lower brake pads 274, 278 are pressed against and engage the wall 268 of the drivable portion 10A. When the brake is released by the actuator 28, the spring 282 is released to allow the fastener 282 and the brake pad to be released. During the positive fortune of the air machine, the brake mechanism 270 is activated at any time to maintain a firm 20 engagement between the upper drivable portion 10A and the lower disposable portion 10B. When necessary, for example, in an emergency, the brake mechanism is released to allow the upper drivable portion 10A and the lower disposable portion 1B to slide relative to each other and become separated. In an aircraft with nested wings, the vacuum system applied to the wings 14A, 14B is released simultaneously with the braking system. In the figure, it shows that the disposable part 10B can be driven up to the front part of the i〇A. 27 200823108 Sliding, but please understand that the disposable part can also slide up behind the driving part i〇a. Figure 43 shows a hydraulic release mechanism 520 that can be used to force an initial separation between the upper body P knife and the lower fuselage portion. The hydraulic actuator 5 (8) and the frame 5 turn #506 can be mounted on the upper body portion 1A, and the lugs 5〇8 are placed on the lower body portion 10B. When the two parts are to be engaged, one of the fishing end 510 of the pivoting arm is fastened to hold the lug 5〇8. When the separation is initiated, the hydraulic _500 pulls the pivot arm 5〇6 in the direction 502. The arm 5〇6 is rotated along a point 504 on the upper fuselage portion and moves in an arched direction 512. The fishing end 51 〇 thus releases the lug 508 and pushes the lug 514 in the direction 514 so that the upper and lower body portions are pushed apart. - When air flows between the upper and lower parts, the two parts are forced to separate further. An alternative method of partial separation is to reverse a vacuum pressure to produce a positive pressure. It is to be understood that the invention disclosed and defined in this specification extends to all of the M.s. combinations of two or more of the individual features mentioned or known in the drawings. All of these combinations constitute different alternative aspects of the invention. BRIEF DESCRIPTION OF THE DRAWINGS FIGS. 1A to 1C are respectively a front view, a side view and a top view of a first embodiment of a modified passenger aircraft of the present invention before separation; 20 FIGS. 2A and 2B are respectively shown in a separated configuration. Front view and side view of the modified passenger aircraft of Figures ^ to ^; Figure 2C shows a top plan view of an upper drivable portion of the aircraft in the separated configuration; Figure 2D shows the aircraft shown in Figure 2C Above the driveable portion of the wing 28 200823108 is a cross-sectional side view along line 2D-2D; Figure 2E shows a top plan view of the lower disposable portion of the modified aircraft in the separated configuration; Figure 3 shows the modified The aircraft is shown in cross-sectional side view along section 5 of line 4A-4A of Figure 1B; Figure 4 shows a detailed cross-sectional side view of one of the releasable engagement mechanisms shown in Figure 3; Part of the figure 3 is a cross-sectional side view in which the upper drivable part and the lower disposable part are in a separated configuration; 10 Figure 5B shows a releasable engagement mechanism of Figure 5A in a separated configuration Details in detail; Figures 6A to 6C are shown separately 4 is a perspective view, a side view, and a cross-sectional elevation view of the releasable engagement mechanism; FIGS. 7A through 7C are perspective views showing the releasable engagement mechanism of FIG. 4 in a 15 release position, Side view and cross-sectional elevation view; Figure 8 shows a detailed cross-sectional perspective view of a seal assembly between the upper and lower portions of the aircraft; Figure 9A shows a wing of the aircraft along the line 9A_9A of Figure lc Cross-sectional side view; 2〇 Figure 9B shows the wing of Figure 9A with the upper and lower wing sections in a separate configuration; Figure 9C shows the vacuum differential to assist in holding the upper and lower wing sections together 10A to 10C are respectively a perspective view, a detailed view and a partial cross-sectional side view of the second embodiment of the releasable engagement mechanism 29 200823108 of the present invention in the engaged position; FIGS. 11A and 11B are respectively shown FIG. 10A is a perspective view and a partial cross-sectional side view of the releasable engagement mechanism in the disengaged position; 5 FIGS. 12A to 12D respectively show a low-height passenger aircraft of another embodiment of the present invention before separation, etc. Angle view, front view, side view and cross-sectional front view, Figures 13A to 13D show isometric, front, side and cross-sectional elevation views of the aircraft of 12A to 12D respectively during separation; 10 14A to 14D shows an isometric view, a front view, a side view and a cross-sectional front view, respectively, of a high-height passenger aircraft according to another embodiment of the present invention before separation, and FIGS. 15A to 15D respectively show the separation of the aircraft of the 14A to 14D. An isometric view, a front view, a side view and a cross-sectional front view of the period; 15 Figures 16A to 16C show an isometric view, a front view, a side view and a side view of a safety facility glider according to another embodiment of the present invention before separation Cross-sectional elevation views; Figures 17A through 17D show isometric, front, and side views, respectively, of the aircraft of Figures 16A through 16C during separation; 20 Figures 18A through 18C show one of the other embodiments of the present invention, respectively Isometric, front view, and side views of a twin-engine aircraft prior to separation; Figures 19A through 19C show isometric, front, and side views of the aircraft during periods 18A through 18C, respectively; 20A through 20C The figures respectively show another of the invention One of the examples is an isometric view, a front view, and a side view of the nested engine high-height aircraft before separation, and the 21A to 21C charts show the isometric and front views of the aircraft of the 20A to 20C during the separation, respectively. And FIGS. 22A to 22D respectively show an isometric view, a front view, a side view and a cross-sectional front view of a high-height launcher of another embodiment of the present invention before separation; The 23D diagram shows the isometric, front, and side views of the emitters of 22A through 22D during separation, respectively; 10 Figure 23E shows a cross-sectional side view of the ARAVS of Figures 22A through 22D during flight into space. 24A to 24D are respectively an isometric view, a front view, a side view, and a cross-sectional front view of the ultrasonic delta wing aircraft of another embodiment of the present invention before separation; 15 Figures 25A to 25D respectively show An isometric view, a front view, a side view, and a cross-sectional front view of the aircraft of 24A to 24D during separation; Figures 26A to 26C respectively show prior to the separation of the transonic delta wing aircraft of another embodiment of the present invention. Corner view, front view And the side views; Figures 27A to 27C show an isometric view, a front view, and a side view, respectively, of the aircraft of the 26A to 26C in the separation period 20; FIGS. 28A to 28C respectively show another embodiment of the present invention. An isometric view, a front view, and a side view of a three-stage aircraft prior to separation; Figures 29A1 through 29C show an isometric view, a front view, and a side view of the aircraft of Figs. 28A through 28C, respectively; 31 200823108 29D to 29G respectively show an isometric view, a front view, and a side view of the upper portion of the aircraft of Figs. 29A1 to 29C of the inorganic tail engine module; FIGS. 30A to 30C respectively show one of the other embodiments of the present invention Isometric, front view, and side view of the biplane prior to separation; 5 Figures 31A through 31C show isometric, front, and side views of the biplanes 30A through 30C, respectively; 32A1 to 32D respectively show an isometric view, a front view, a side view, and a cross-sectional front view of a double-wing high-altitude transmitter of another embodiment of the present invention before separation; 10 Figures 33A to 33E show the first 32A1 to 32D transmitters during separation Angle view, front view, side view, and cross-sectional front view; Figures 34A to 34C show an isometric view, front view of a crank-V aircraft prior to separation, respectively, according to another embodiment of the present invention And the side view; 15 Figures 35A to 35C show an isometric view, a front view, and a side view, respectively, of the aircraft of the 34A to 34C during the separation; FIGS. 36A1 to 36C respectively show one of the other embodiments of the present invention The isometric, front and side views of the box-wing aircraft before separation; Figures 37 to 37c show the isometric, front and side views of the 36ak36C aircraft during the separation period of 20 Figure 38A to 38D are respectively an isometric view, a front view, a side view and a cross-sectional front view of a high-height passenger aircraft incorporating a slide-release mechanism in accordance with another embodiment of the present invention before separation; And Figure 38F shows an enlarged cross-sectional elevation view of the 38B to 38D barrier system before the separation 32 200823108; Figures 39A to 39C show the isometric, side and transverse views of the aircraft of the 38A to 38D during separation, respectively. Sectional view; Figures 40A to 40C show the 38A to 38D, respectively Isometric, side, and side views of the machine after separation 5; Figures 40D and 40E show enlarged cross-sectional elevation views of the 38A to 38D barrier system after separation; Figure 41 shows sections 38A to 38D Amplified cross-sectional elevation view of the aircraft prior to separation with the brake system actuated; 10 Figures 42A-42B show enlarged cross-sectional elevation views of the actuated braking system of the aircraft of 38A-38D during flight; The figure shows a hydraulic release mechanism for separating the upper and lower portions of the aircraft. [Description of main component symbols] 10... Modified passenger aircraft 22... Sealable interface 10A··· Upper driving part 24... Passenger compartment 10B··· Lower part 26... Main cockpit 10C···Tail module 27...Auxiliary driving area 12...Body 28...Exit door 14A...Upper wing part 3〇...Front and rear cargo compartment 14B...Front wing part 32,34··· Front and main landing gear compartments 16,18·· Wing-mounted engine 35, 36················································································ ...inflating chamber 48...upper ceiling 50...releasable engagement mechanism 52...housing 54··lower flange 56,112,114,500~hydraulic actuator 58...claw assembly 60...expanding head 62...release bolt 64 ...frame 64A, 64B...first and second separate triangular plates 66... center pin 68, 70··· inner and outer jaws 72...restricted sockets 74,76,116,118~lever arm 78...hydraulic hammer 80... Cylinder block 84... outer lever arm 88... inner pole shaft point 90... horseshoe hook assembly 92... cable or coupling rod 94... sealing gasket 96... Lower floor 98...film material 100...fuel tank 102···vacuum manifold 110...releasable joint mechanism (fig.10A, 10B), vacuum manifold (fig.9A, 9B) 102···clear side cover 104 ···vacuum manifold arrangement 106...path 108···rib 120···rotating tweezers 122...arching stop 124···cylindrical cavity 126...socket box 200···v type hull 202 ...spray strip 204··· can deploy the ventral horizontal stabilizer and elevator 206... fairing 208... support 210···air intake 212...horizontal stabilizer 34 200823108 220...low earth orbit re-entry vehicle Transmitter 266···Wall 220···ΑAutonomous rocket-assisted carrier 267...Opening section (ARAVS) 268...Lower horizontal wall 220Β···Reusable high-altitude transport 269...seal transmitting transmitter 270...Brake mechanism 221...payload 272...upper grooved brake pad 222...liquid (or refillable solid) fuel 274,278...lower brake pad 226···tail panel 276...upper brake 塾228···wing lock Fixing mechanism 280...Brake actuator 230...pillar 282...fastener 232···surface landing stabilizer 284...compression spring 250···Sliding system 502...Direction 252...Rack 504···Point 254...Gear 506...Pivoting arm 256···Actuator 508,514...Lug 258···Upper grooved roller 510...fishing End 260, 264 ··· lower roller 512... arch direction 262... upper roller 520... hydraulic release mechanism 35
Claims (1)
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| PCT/AU2006/001247 WO2007022601A1 (en) | 2005-08-26 | 2006-08-28 | Aircraft safety system |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| TW200823108A true TW200823108A (en) | 2008-06-01 |
Family
ID=44802528
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| TW095143198A TW200823108A (en) | 2006-08-28 | 2006-11-22 | Aircraft safety system |
Country Status (1)
| Country | Link |
|---|---|
| TW (1) | TW200823108A (en) |
-
2006
- 2006-11-22 TW TW095143198A patent/TW200823108A/en unknown
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