發展與驗證一具高度模組化與高精準度的發射載具導控模擬器

Abstract

一般來說，火箭從發射場垂直起飛，穿越大氣層後，藉由自身的推進器在太空中加速至可繞行地球的慣性速度後，使酬載入軌繞行地球飛行；在這整個飛行過程中，會有一些不可預期的擾動出現，例如額外的氣動力影響、陣風影響、感測器本身的誤差與致動器實際作動和指令的誤差，導航的演算法跟飛行的時序將大大的影響火箭最後入軌的精確度，所以需要去發展一個工具可以考慮上述的擾動、導航演算法與飛行程序對於入軌精確度的影響。 本論文前半部份修改現有公開的導控程式成為一高度模組化的模擬程式, 同時進行一系列與原本導控程式模擬結果進行比較, 確定修改後程式無誤。後半部份即利用新完成的模擬器進行模擬一假想火箭入軌的飛行過程。模擬器考慮前述的擾動、導航控制與程時序控制的影響。飛行程序依序包括：火箭起飛後使用TVC使火箭傾斜一角度，第一節引擎熄火、拋棄，第二節火箭開始進行重力轉向而為了將氣動力損失降至最低同使也控制火箭使其攻角接近0度，當第二節引擎熄火後，火箭會開始滑行一段的時間，當姿態到達一定的航行角(Flight Path Angle)後第三節引擎啟動，開始進行Powered Explicit Guidance (PEG)閉迴路導航，在這個階段火箭使用RCS或TVC控制姿態進行入軌。 模擬結果顯示在適當的航行角下開起PEG控制入軌將會對最後的入軌精度有很大的幫助，更進一步將探討不同的航行角下開始進行入軌控制的入軌精度。本模擬器的完成對日後進行 Processor-In-the-Loop (PIL) 與Hardware-In-the-Loop (HIL)的研發將有相當重要的助益。

In general, satellite launcher takes off vertically from launch pad, flies through the atmosphere so that it can perform orbit insertion where rocket shall propel itself up to the designated inertia velocity required for its payload to orbit the Earth. During the course of flight, several disturbances, such as aerodynamic force, wind gust, imperfection of sensor/actuator between real and measurement/command, guidance law, autopilot law and timing of flight sequences will ultimately impact the accuracy of final orbit insertion. As a result, there is a need to build a simulation tool to evaluate the impact toward the accuracy of orbit insertion owing to those disturbances, laws and timing of actions during the flight. In this thesis, a highly modularized simulation tool with high fidelity capable of simulating the flight of a rocket based on its physical properties is modified based on publicly available GNC (guidance navigation and control). The newly developed code has gone through numerous validation simulations against the original code. The simulator considers these aforementioned disturbances, navigation, guidance and control laws and sequence of ignition/separation actions along with designated timing. The sequence of actions during flight are summarized as follows. First, thruster vector control (TVC) vector tilts at an angle immediately after rocket taking off. After the jettison of first stage, the second stage starts a gravity turn in a near zero angle of attack manner to minimize the aerodynamic loss during flight. After the second stage engine cuts off, the rocket coasts for a certain period of time until its attitude reaches a appropriate flight path angle. Later, the third stage engine ignites and the activates a close loop control using reaction control system (RCS) or TVC with guidance using Powered Explicit Guidance (PEG) will be enabled for accurate orbit insertion. For the proposed flight sequence, the simulation suggests that the timing or the proper flight path angle to enable the close-loop PEG on the third stage will impact the accuracy of the final orbit insertion. Further investigation of this particular factor is performed by the comparison of orbit insertion accuracy for different flight path angles. Completion of this simulation tool shall benefit our future development in Processor-In-the-Loop (PIL) and Hardware-In-the-Loop (HIL).