火箭燃料用环氧树脂系列粘合剂的工艺特点对比

火箭燃料用环氧树脂系列粘合剂的对比研究

科创航天局301所

目前局内使用环氧树脂作为火箭燃料粘合剂的趋势日益递增,相信广大网友对于这种常见的粘合剂并不陌生。 但由于牌号不同,固化剂不同,所导致的粘合效果以及燃料的力学性质差异性是显而易见的 近日,我在KCSA-301中进行了较为简单粗略的对比性实验,现将这些不同告知广大爱好者 以便在以后固体燃料的制造过程中,避免不必要的麻烦。

首先,我们在进行对比性实验之前,先要对一些牌号所代表的物质及意义进行简要阐述。

1.环氧树脂的牌号数字 例如:E51 即表示其环氧值为51 环氧值是指每100克环氧树脂中含有的环氧基的当量数。单位为当量/100克。它与环氧当量的关系为环氧值=100/环氧当量。

E57,E51,E44都属于双酚A型环氧树脂 因此符合以下换算关系: 环氧值=2×100/环氧树脂分子量

即Ev=2×100/M 环氧当量=100/环氧值 即En=100/Ev

环氧基含量=43×100/环氧当量 即Ec=43×100/En

不需要复杂的计算便可得知,同一类型(例如:双酚A型)的环氧树脂的数字牌号越大,其聚合的分子链对应的分子量则越小, 其树脂的流动性越好,相对的反应基团的屏蔽效应也较小,反应活性较高,速率较快。但由于链长相对较短,其使用同一种方式聚合固化的产物力学性质也会较差(表现在抗拉强度上,俗称“较脆”) 。

2.固化剂所对应的固化类型 常用固化剂有T31及GCC137 T31属于多乙烯多胺类固化剂(粘度高,密度大,反应活性高)。 GCC137则属于聚醚胺,异佛尔酮二胺类固化剂(粘度低,密度较低,反应活性低)

3.增塑剂的类别 DOS,DOP,DOZ都属于高级烃的之类化合物,在增塑体系中,并不参与反应,只能部分溶入体系中,从而达到增塑效果 HTPB在环氧树脂体系中,属于特殊性增塑剂,其HTPB的端羟基会与环氧树脂的端位环氧基团进行化合反应,从而达到链增长的增塑效果。

以下是對比性實驗記錄:

實驗前,準備好試管及各種原料

將E51及E57分別加入到特定的試管中(每根試管中3.4g)

每根試管中3.4g

加入HTPB到特定的試管中(每根試管中加入1g)

分別將T31及GCC137加入到對應的試管中(每根试管1g)

將其迅速攪拌至均勻

然後放入真空倉內進行脫氣處理(真空壓力≤-0.09MPa)

取出後發現E57和T31的那一組已經固化了,放熱巨大,手觸感覺後,估計已經超過60℃)
攪拌後,進行到此步的時間大約為15min,室溫28.5℃
分段是由於抽真空脫氣時底部氣體膨脹所導致

斜過來45°對比一下

以上實驗均在室溫28~30℃的環境下進行

得出的結論是:

以混合后开始计算时间

E51-T31,E57-T31,E51-T31-HTPB,E57-T31-HTPB这四组以T31作为固化剂的体系中
在混合20min后均有强烈放热,手触估计温度>60℃。放热的强烈顺序E57 > E57-HTPB > E51 > E51-HTPB。
约40min后,体系完全固化。
E57-T31和E51-T31在没有良好的降温措施下,会产生大量气泡!!

E57-GCC137,E51-GCC137,E57-GCC137-HTPB,E51-GCC137-HTPB这四组以GCC137作为固化剂的体系。
在混合20min后,粘度略有上升,并且不明显,其体系仍然有很好的流动性,可以顺利进行加工。
在混合5h后,E57体系完全固化。
在混合6h后,E51体系完全固化。

从以上现象中,我们可以总结出GCC137是一种活性较低,利于延长固化时间,利于药柱加工的一种固化剂。

而E51体系的粘度普遍高于E57体系。

另外还需说明的一点,在以上反应中,混有HTPB作为增韧剂的体系均产生了少量分层。

这表明HTPB的用量超过了粘合剂所能反应的范围。

这是前一天晚上做的60℃环境固化实验结果。
使用E57及E51与T31及GCC137进行对比。
结果惨不忍睹~~~~

以E57作为粘合剂的体系迅速反应,放出大量热量,以至于体系迅速膨胀,产生大量气泡。

背面

E57-GCC137氣泡特寫

E51-T31体系在60℃环境下固化同样剧烈放热,产生了气泡,但是不是特别明显,所以没有拍照记录。

综上所述

可得出一种较为使用于固体燃料的粘合体系配比

即 E57                   74%
GCC137            21%
HTPB                   5%

先将HTPB与E57预先混合至乳白色,反应10min后,再与GCC137混合进行固化。

其优点是,粘合体系固化前粘度较低,加工性质良好,固化时间长,固化后仍能保持良好额力学性质以保证大型药柱的机械加工需求。(撰文:胡振宇)

液体火箭发动机设计参数计算软件发布

科创航天局贵州局的爱好者为设计液体火箭发动机编写了一系列计算工具,现在发布其中一部分,由魏广寅编写。

软件合集下载:LREDesignTools

1、气体直流喷注器设计

该软件系根据热力学公式,利用VB编写,考虑了气体可压缩性,以及管道中流动的内能-动能(温度-动量)转化,能设计出与实验结果十分接近的参数。

软件使用说明:气体比热比、气体常数、气体初始温度等可根据我之前发布的PEP软件计算得出,初始压强与背压之差即为喷注压降,质量流量根据发动机内弹道软件计算得出,流量系数对于普通光滑圆管(雷诺数>2000)的结构,当管道长径比为3左右时,可取0.75-0.85。喷口数量根据喷注器类型和钻孔工艺自行设定。

2、液体直流喷注器设计

该软件系根据流体伯努利方程,利用VB编写。适用于粘度较低的液体喷注计算(例如煤油、汽油、乙醇、甲醇、液氢、肼等)。

软件使用说明:液体流量根据内弹道设计得出,液体密度查表得出,喷注压降和喷嘴数量根据喷注器形式和雾化要求自行设定。流量系数一般取0.4来进行初步设计,之后根据实验对该系数进行修正。

3、气蚀文氏管设计

气蚀文氏管主要用于稳定管路的流量,其流量只决定于上游滞止压强和液体在当时温度的饱和蒸汽压,其流量不随下游背压的变化而变化。可有效隔离燃烧室燃烧震荡和粗糙燃烧对燃料管路及储罐的反馈。

该软件系根据液体伯努利方程及气蚀条件方程,采用VB编写。

软件使用说明  :
流量由发动机内弹道参数得出,饱和蒸汽压查表得出,管端直径即为管路直径。

75毫米RNX固体火箭发动机的制造和试车

科创航天局第301研究所承担的WH-4型探空火箭项目,目前已经进入了最后的冲刺阶段,争取每周一帖的进度……

今日(2012年6月2日)下午五时许,KCSA-301的四名成员前往“沙地”测试厂进行75规格RNX发动机试车,成果喜人~~

该台发动机,从设计到制造花费了一个多月的时间,其中从结构设计,到各类零件的加工都下足了功夫。

好,不说其它的了,现在就来对此次活动的历程进行介绍……

首先在总结之前的发动机事故和不足的基础上,我们认为在发动机的燃气密封上应该狠下功夫

因此,这次在设计的时候,特意在主要转接处进行了硅橡胶多层密封的设计

(白色为PVC隔热壳体,红色为金属连接处的梯形丝牙,黄色则是硅橡胶密封圈及密封垫 )

为此,我们还特意使用CNC对星孔药柱端面的硅橡胶密封片进行了开模制造,以下为粗模

之後進行220目油石打磨

400目粗打磨
  

由於,矽膠片的表面精細程度要求不需要太高,因此就沒有進行鏡面處理,直接用數顯銑床進行插銷位鑽孔
  

最後,在200℃高壓硫化床進行高溫硫化操作,成型脫模

最終的產品,相當和諧~~

之前製作出來的RNX藥柱,這次用的是E57和GCC137為粘合體系,少量HTPB進行力學改性,配以車床和CNC做出的星孔藥柱

4根備用,每根規格都是200長度
  

嘎嘎~   今天中午,備齊所有裝備

发动机数据如下:

壳体材质:40Cr
喷口堵头材质:45#
喷吼材质:热解石墨(可千万别小看这东西强度,简直要命)
有效燃烧室长度 600mm(不包含密封件厚度,单纯燃料长度总和)
发动机内径:73.5mm
发动机外径:86 mm
设计耐压:98.1MPa (初次测试,就高些了,以后等数据出来就可以慢慢车薄)

数据采集设备:

KCSA-301 V1.1版高速数据采集卡(由罗澍设计制造)
KCSA-301 T-20125 中型固机试车平台(由胡振宇设计制造)
中航L6G 750kg行程应变片  200Hz采集频率 (感谢刘虎支持)
PTB503  32MPa  1000Hz 采集频率

开始组装~~

焓熵说了“药柱啥的,大朗头伺候”

  

嗯嗯,好的很~沒有偏心
  

來一張RNX菊花圖~~ 話說挺不錯的

這東西實在太重了,還是上卡盤解決問題

不行! 還是要水管扳手……
  

測測重量…………  我勒個去,嚇死人~~14.9kg

出發~  咱們去“沙地”測試廠……   可憐的我,只能人貨混運了
  

好大一包銀藥啊~~

會會牌,冷卻系統
  

最後的全景唉~~
  

測試後的噴口,感覺還過得去
  

測試裝置全圖

噴口擴展段特性,有不少碳沉積物

Warmonkey倒出推壓力曲線

试车视频:

http://v.youku.com/v_show/id_XNDA2ODMxODU2.html

 

最终结论:

此台75规格RNX发动机,工作时间越6s
总冲2554.0436N*s
比冲由于药柱质量暂不确定,以后补上,粗略估计的值有些低,只有90.6708s
峰值推力>100kg
点火压力:2MPa

发动机工作时间较长,喷燃比略低,并且有部分时间处于尾焰无推力状态,失压工作

改进办法:
1.提高设计喷燃比
2.缩小喷口喉部直径

设计喷燃比曲线

後續還有噴口灼噬報導,現在開閘放水~~

晚上回到KCSA-301之後,俺和猴子同學進行了發動機屍檢工作~

噴口打開,發現沉積物很多

石墨啊石墨~你還健在啊啊啊…………  這個情況很好,打消了老虎的顧慮
我不得不說,石墨這個東西太難拆下來了,最後只能全部敲碎,才能一小塊一小塊的翹下來

噴口收斂段與石墨轉角處,石墨的弧面被沖刷成平面了

菊花無處不在~

用砂紙清洗完之後,發現收斂段的金屬被沖刷了不少,出現了星孔形狀的凹陷
  

堵頭處,密封圈完好,表面略有沉積物

  

殼體內部完好,靠近堵頭處有沉積物

內部略有微量沉積物

清洗噴口收斂段的時候發現了很嚇人的一幕,由於沖刷嚴重,在噴吼處,金屬沉積到了石墨上,
形成了薄薄的一層金屬層,這層沉積物一直延續到擴展段轉角處

慘不忍睹的隔熱層…………歸其原因,是因為發動機內壁無法加工,只有73.5mm內徑,所以車去了不少PVC,導致最後的隔熱層厚度只有1mm

结语:此次发动机测试总体来说相当成功,但细节设计上仍有需要改进的地方

首先,从燃烧而言,此次发动机的工作压力较RNX常用的工作压力略低,这是由于设计时,喷燃比曲线设计留下的。
另外,在设计过程中没有中和材料原料规格的因素,因此出现了隔热层过薄的情况,在以后的大规格发动机中,将采用环氧树脂复合材料进行刷涂隔热
除此之外,在材质的选择上,45#已经难以胜任喷口这一零件,在以后的制造中,均将换成40Cr材质制造喷口,并进行热处理
在喷口收敛段和扩展段上,此次在收敛段采取了抛光处理,但由于时间因素,此项环节仍然没有做到最好,还是略有车加工时留下的细微刀纹,在以后的制造中,均采用200目,400目,800目,1200目,最后使用抛光砂纸进行全面打磨,以消除刀纹对气流的影响

最后来一张会会的邪恶点火器~

  

全文完。

YT-3发动机爆炸原因分析总结(1)金属结构

作者:刘虎

万户一号第一级装备的是YT-3发动机。该发动机试验机共生产四套,工作五次(其中一枚重复工作一次),有二次成功工作的记录,三次失事爆炸的记录。爆炸原因是多方面的,既有燃料燃速超出预期,也包括发动机耐压不达预期。本帖仅分析发动机金属结构的原因。

YT-3发动机为内径50mm,壁厚2.55mm的铝合金管式发动机。两端承压结构用45号钢车制,采用径向螺栓连接,螺栓12颗均匀分布,螺栓规格4mm,丝扣长约2.5扣。

壳体铝合金标称牌号6061,T6。
密度:
屈服强度:
从断口情况看,该铝管的质量不佳,能否达到6061T6的标称强度,有待检验。

破坏情况:

裂口在靠近头部一端最大,破坏最明显。12.18试验,头部完全破坏;10.17试验,头部裂开,承压结构未飞出。两次试验,均在靠近头部的三分之一位置处破坏最大。

查见螺栓连接处有明显滑动或形变。头部见铝管圆孔变形,沿轴向拉成椭圆。尾部(喷管)见喷管承压结构与铝管错动,向外到达螺栓活动极限位置,圆孔轻微变形。

头部承压结构,螺栓孔折弯,部分丝扣破坏或被拉光。

12.18试验发动机的大裂口明显呈不同的两层裂痕,一层光滑,另一层粗糙。另一侧裂口内显示显著瑕疵,瑕疵靠外为断裂痕,靠内为整齐界面,有氧化物附着。综合判断,两处应均为原有瑕疵。

10.17试验,裂口处未发现焊缝,但是裂口最大处有明显拉薄现象,有两处长30mm左右的Y字型断面呈刀口状,外侧有灼热氧化痕迹,内侧有高温形变痕迹,为最初破坏点。发动机爆炸后飞出,撞击发射架,在该处留下切痕,本应具金属光泽,但现已氧化变色,说明爆炸时温度很高,应为高温失稳破坏。两处最初破坏点的Y字形中点距离两端的分别为145mm和143mm,正好位于药柱分段接头处,证明上述判断正确。裂口两端为撕裂状,判断螺栓处系沿大裂口撕裂,不是最初破坏点。未破坏的部分,见一条内外侧均有的痕迹。

计算(拔刀斋):

按6061-T6管材的技术数据,算得该铝合金管的理论破坏压强为25MPa。

按屈服强度数据,得螺栓压强达270MPa时,螺栓处铝合金发生破坏。按此计算该发动机极限耐压为13.7MPa。

当发动机内部压力为13.7MPa时,每个螺栓受剪力为:2252N,约230公斤力,螺栓不会被剪断。

讨论:

发动机设计工作压力为6MPa,指喷管收缩段起始处的压力。实际头部压力应高于6MPa。按工作压力5倍选取耐压值,试验压力应达到30MPa。铝管的理论耐压能力与之接近,而螺栓处成为薄弱点,残骸可见螺栓处破坏较大,也证明了这一点。
如果管材质量尚好,更可能发生的情况是头部剪切飞出,铝管基本完好。事实于此相悖,高度怀疑管材质量问题,从显著瑕疵可以判断,该管材为劣质管材。
发动机点火后温度会上升,铝合金外壳强度会下降。隔热保温措施是否充分,以及工作温度下的强度是否满足要求,均应检验。

结论:
12.18试验
1、发动机金属结构设计有误,螺栓太细,丝扣太短。
2、劣质管材。
3、缺少试验项目(水压试验)。

10.17试验
1、发动机金属结构设计有误,螺栓太细,丝扣太短。
2、耐热不足。(应为隔热不足,但在金属结构这个角度看,就是耐热不足了)
3、疑为劣质管材。
4、缺少试验项目(水压试验)。

两者皆有
1、第一次试验爆炸后,未彻查原因。
2、装药方式有误。铝管采用分段装药,应当外加隔热层保护分段处,而两次均未加装。

12.6试验发动机残骸

12.18试验发动机残骸

发动机喷管,其它试验残骸,头部(12.18)、尾部(12.6)印模。

火箭平台上引力蓝移现象的观测(英文)

本文受到学术造假质疑,KCSA可能从网站上撤销本文的发表,详情请见:http://bbs.kechuang.org/read-kc-tid-41784.html
火箭平台上引力蓝移现象的观测(英文)
李雨翀 广东广州
1. Abstract
An experimental verification of general relativity conception was made using an atomic maser in a rocket attaining an altitude of 1,800 metres.
The signal of the maser was monitored on the ground, so that the effect of gravitational potential on the frequency of the maser was measured.
The resulting data was processed through a careful prediction and elimination of the Doppler shift and other error resources, so that the gravitational blue shift is directly observed. The experiment is described including a consummate discussion of navigation algorithm applied in the processing procedure. The authors believe that this is a direct high accuracy test of the general relativistic phenomena using an airborne clock.
2.Introduction
A rocket is constructed and launched carrying an atomic frequency standard as the payload. The frequency of the signal from the atomic frequency standard is examined on the ground.
Another atomic frequency standard is used as a comparison to monitor the change in frequency of the received signal from the payload.
A group of variables that will influence the change are eliminated so that the resulting data, representing the relativistic shifts, are recovered andrecorded.
The objective of the experiment is to test general relativity concept by measuring directly the effect of gravitational potential on thefrequency of a proper clock, in this case the atomic frequency standard.
In this experiment, a gravitational effect amounting to 5.6e-13 was measured.

The predicted proportion change in frequency is expressed in equation [a].Where β is the velocity/c and r is the displacement of the rocket relative to theground base. ais the centrifugal acceleration of the ground station while ε represents thepropagation vector of the rocket-to-ground signal. Our knowledge of the relative velocityand displacement of the rocket is obtained from the flight data recorder installed in thepayload. As we chose the ground base as the navigation frame, the movement of the groundbase in the geographic frame was straightly eliminated in the
navigation system.
In equation [1], the first term is the gravitational blue shift, the second term expressesthe Doppler shift. The last term describes the effect of the rotation of the earth during thepropagation of the signal. In the elimination of the second and third terms, out knowledge ofthe rocket’s velocity and position are obtained from the FDR(Flight Data Recorder) while theknowledge of the velocity and position of the ground base is gained from the Earth Modeland GLarLng of the launching site.
The specific procedure is exhibited in the materials and methods section.

3. Materials and Methods

The data was first generated and collected after the hardware processing of ground base, which procedure is labeled as data
acquisition. The recorded data was then stored in the computer for further processing, labeled as Data processing.
Data acquisition
a) Algebra description
The frequency signal transmitted from the payload is fixed at f0 + △f.
△f was set to be 50mHz (Figure [d]).
The signal received on the ground base is labeled as f1. It is predicted to be

after the elimination of errors.
The ground station processed f1 with a standard signal of frequency f0 , which was generated by the atomic frequency standard.
Heterodyne-beat method was applied thus a signal with frequency f1 – f0  was sampled by a high speed analog to digital converter as f2 and processed by the digital signal processor. A standard signal of frequency△f is generated by atomic frequency standard and processed in the computer, a series of effects including doppler shift are taken into account.

The final signal, labeled △fR, is expressed by equation [2].
The frequency of f2 was estimated by 4-parameter estimation algorithm and recorded in the computer. The method of 4-parameter
estimation will be discussed later.

Therefore, the resulting data was obtained. The predicted behaviour of this final signal is shown in figure [6].

b) Technical details
A transmitter and a superheterodyne receiver were built specially for the experiment.
The signal from the on-board AFS(Atomic Frequency Standard) was directly amplified and transmitted. The output power of the amplifier was +38dBm and the frequency was 63.8978MHz.
The structure of the receiver is shown below.

The signal from the antenna at the ground base was filtered and directly amplified by a LNA(Low Noise Amplifier) and mixed with FS1 by mixer 1 to obtain an IF(Intermediate Frequency) with a frequency of 5MHz. The IF signal was processed by an AGC(Automatic Gain Control) circuit so that the amplitude was stabilized. This processed IF signal was mixed with FS2 by mixer 2 to obtain a 50mHz signal and was converted into digital signal by a high speed ADC. The exact frequency of the signal sampled by the ADC was estimated through 4-parameter estimation algorithm. FS1 and FS2 were generated by the AFS at the ground base.

1) The realization of heterodyne-beat method
The kernel of heterodyne-beat method is the shift of spectrum. As a result, a mixer is used to obtain the difference in frequency of two signals. In this application, two integrated mixer circuits AD831 were used.
The following images indicate a test result for the mixer circuit. The left image indicates the signal of RF input of the mixer and another one shows the local oscillation input of the mixer. They were generated by two DDS(Direct Digital Synthesis) circuits. The clock standard of the DDS circuits was connected to an AFS. The frequencies of them were 7MHz and 7.00001MHz.

The following image shows the output of the mixer which equals to the difference between two frequency signals connected to the local

oscillation input and the RF input. Here, the difference was 10Hz.

2) The design of AGC circuit

An AGC circuit was used to stabilize the amplitude of IF signal.
The input signal was demodulated and filtered into a voltage signal represents the strength of the signal. This voltage signal was used to

control the gain of a VGA(Variable Gain Amplifier).

In this case, the wave-detector was AD8307 and the VGA was AD603.
A test for the AGC module is shown below. Channel 1 was connected to the output of the AGC module, channel 2 was connected to the input. The input of the AGC module was connected to a function generator.

As the waveform shown on the oscilloscope, although the amplitude of the input signal to the AGC was changed, the amplitude of output signal of it remained the same.
A test result for the transmitter and the receiver is shown below.
The transmitter was placed 800 metres away from the receiver. The image on the left shows the output signal from the transmitter and another image shows the signal sampled from the output of LNA of the receiver.

The image below shows the signal sampled from the output of the AGC module, where frequency = IF = 5MHz.


2) Method of estimation of frequency
a) Definition of 4 parameters Assume the sampled signal S(t) is given by

Where A0′ is the ideal amplitude of the signal, ω0′ is the ideal frequency of the signal, C0′ is the ideal DC offset of the signal and θ0′ is the ideal phase of the signal.
The signal can be expressed by the equation

Where

Suppose the magnitude of the signal sampled during time tk(k = 0, 1, 2, …) is y(k), is given by

b) Method of 3-parameter estimation
Suppose the sampled voltage value of the signal at time tk is y(k), k = 1, 2, 3 …N-1. The amplitude of sine, amplitude of cosine and DC offset is defined as A,B and C. The RSS(Residual Sum of Squares) between the estimation value and actual value is given by

Where N is the length of samples, set

The solution for X is given by the least square solution below:
.  
c) 4-parameter estimation algorithm
The idea of successive approximation is applied in this algorithm. First, a rough frequency is given, 3-parameter estimation algorithm is applied to the sampled signal. The cosine amplitude, sine amplitude, DC offset and estimated RSS are obtained. The operation is repeated with
different frequencies so a serial of estimated RSS are obtained. One of those set of obtained result with minimum estimated RSS is the value of actual frequency. The detailed steps are shown below.
1) Determine the frequency of the signal roughly though DFT(Discrete Fourier Transform), label this frequency as f0.
2) Set the domain of iteration to be ωdl and ωdu, where ωdl is the lower boundary, given by ωdl = f0 – fclk / N. ωdu is the upper boundary, given by ωdu = f0 + fclk / N. fclk is the frequency of sampling clock and N is the length of DFT.
3) Set ω0 = ωdu – ωdl. 2M+1 points (M∈N*) are samples between ωdl and ωdu with equal intervals. 3-parameter estimation algorithm is used here to compute the RSS of this group of samples.
4) Find and record the minimum value of RSS of samples in step 3 This minimum value is corresponding to the actual frequency.
Repeat operations 2 to 4 until the precision of the estimation reaches the required level.
A picture of the receiver is shown below.

2) Data processing

a) Downlink signal

The navigation system provided data of dynamics with a sample rate of 1,200 samples per second. The data was given in terms of
angular velocity and acceleration in the on-board coordinate. The six groups of parameters are ωx, ωy, ωz, ax, ay and a–z, respectively where the X-axis if the mean axis of the rocket.
In the determination of change in angle, △θ and change in velocity, △V, the cubic spline function is adopted in curve fitting before integration. This method of Simpson’s rule provides six groups of data: △θx, △θy, △θz, △Vx, △Vy and △Vz. As a result, the behaviour of the rocket between samples are predicted and considered.

In the determination of attitude angle, the method of Quaternion is applied.
The quaternion numbers at time tm+1 are given in equation [9], where △θx, △θy and △θz are the output of change in angle and vector Φ

is the rotation vector, which is given by equation [9]

In equation [9] and [10], the angular velocity of the rocket is assumed to fit cubic function. However,  the actual angular velocity does not fit a cubic function.
Equations [9] and [10] do not achieve minimum shift of algorithm. After the parachute deployment, the rocket was suspended in the descending stage. Thus the rocket is likely to experience coning motion, which means that the rocket vibrate about the equilibrium position at small angles. The coning motion is the worst working environment for the SINS(Strapdown Inertial Navigation System) as it will cause serve shift of the Math Platform.

For optimization algorithms, the following improvements are made.
O-XYZ represents the reference frame R, which is the on-board frame when the rocket is in equilibrium.

Let b(tm-1) and b(tm) to be the instantaneous on-board frame at time tm-1 and tm.
According to Euler Theorem, O-XYZ can be regarded as a rotating transformation of b(tm) or b(tm-1) with rotating vector Q(t) given in equation [11].

The shift on-board frame can be regarded as a rotation transformation of the ideal on-board frame, which is the frame when the rocket is in equilibrium. The samples of △θ is grouped in three again.
In each group, the samples are labeled as .

Equation [12] is an improved  form of equation [3]. By selecting proper constant k1 and k2, the effect of the coning motion is minimised.
Here, the ideal values for k1 and k2 are 0.45 and 0.675.
Therefore, the attitude of the rocket is found through the optimized quaternion algorithm. The Eular angles are found by equation [13].

The quaternions are supposed to be standardized. However, resulting form calculation errors and other factors, the quaternion numbers gradually loses standability. The standardization of quaternion numbers is applied at the end of each period of attitude refreshment. The formula for standardization is given by equation [8].

Where  is the standardization value and  is the value after attitude refreshment.
So far, the discussion of the rocket’s dynamic is in the on-board coordinate. However, the final results have to be expressed in the navigation frame, which sets the ground base as the origin.

Equation [15] gives the coordinate transformation matrix (attitude matrix) from on-board frame to navigation frame in terms of quaternion.
The initial extraction quaternion numbers are thus given by the initial attitude matrix obtained in initial azimuth alignment.

The velocity of the rocket at time tm in the navigation frame, Vm, is given by equation [16], containing a series of error compensations.
Vm-1 is the velocity in the same frame at time tm-1. Cm-1 is the coordinate transformation matrix at time tm-1. Vm-1 is the compensation velocity caused by while△Vg/corm is the compensation velocity caused by the deleterious acceleration. △Vsfm is the compensation velocity caused by ecific force.

Where△Vm us the change in velocity during period [tm-1, tm].
△Vrotm is the compensation velocity caused by rotation effect.
△Vsculm is the compensation velocity caused by sculling motion.

Due to air current, gustiness and other factors, the rocket experiences vibrations during the flight. Those factors cause a highly dynamic working environment for the payload. Therefore, the velocity has to be compensated so that the sculling effect and the rotation effect are
eliminated. Otherwise, the calculation of velocity will involve serve errors. When it comes to position determination, there two error resources contribute to scroll errors. Here△Vrotm and△Vsculm represent the compensation velocities due to the rotation effect and the sculling effect
respectively. the rotation effect happens when the direction of linear velocity rotates in a three-dimensional coordinate. The sculling effect is caused by the angular vibration and linear vibration are in phase and of same frequency on the rocket. This is quite similar to the sculling motion: on one hand, the syrup vibrates periodically about the lateral axis of the boat. On the other hand, the boat forges ahead along the direct-axis in an intermittent behaviour.

The original expression for △Vrotm and △Vsculm are:

The optimized formula for △Vrotm and △Vsculm are

The optimized algorithm for sculling effect rotation effect, in velocity determination as well as the correction for conning motion in attitude determination make sure that the motion of the layload is precisely calculated in spite of the unstable motion of the rocket. Hence the cancelling of doppler effect and second-order general relativity conception shift are more reliable. The specific precision level is related in the discussion section.

The final expression for Vm is given by equation [22]:

Considering that all the compensation dosages have been taken into account the calculation of Vm and that the data is discrete with equal time internals, the data of displacement, is obtained through numerical integration.
As the dynamic data is determined, the following equations are substituted into
equation [2].

Where c is the velocity of light, Wen is the angular velocity of the earth in the navigation frame and  is the radius vector of the earth at the launching spot.
The frequency of this signal is plotted against time in Figure [5].
The navigation algorithm aims to calculate the velocity of the rocket in the navigation frame. However, the accelerometer does not tell deleterious acceleration and relative acceleration of the rocket. Therefore, the compensation velocity has to be estimated from the measure value.

Where g is the gravitational acceleration of the launch site.
In equation [24], the second term represents the centripetal force of the navigation frame, which rotates about the earth. The third term is the coloris acceleration due to the interference of the  and . The coloris acceleration is when the rocket experiences a relative velocity to the navigation frame, while the navigation frame rotates itself.
Substituting the data obtained from the FDR module, which are ωx, ωy, ωz, ax, ay and az during the flight into the equation [24], data of
velocity is obtained.
A compensated with the data of frequency monitored from the ground base, the dynamic data is applied in the following Doppler-
cancelling system.
The doppler shift is given by equation [25]

Theta is given by

In the former process, the dynamic data of each sample is recorded with its corresponding frequency monitored.
Thus doppler shift effect of  the downlink signal eliminated. Now this signal is sipposed to be given by the following equation

This is the pure relativistic shift of the downlink signal.

b) Comparison group

The data of the transmitter on the payload, was real time recorded, which is linearly related to the time standard on the payload. As the payload experiences a relative velocity to the navigation frame, the clock effect is considered. The following cancellation applies to the special Relativity Conception. For the time base, the original time interval between two pulses tm-1 and tm is △    t.

The data of the time base experienced former processing with the dynamic data, thus the actual transmitted signal is calculated from the corrected time base in the computer. This signal is labeled as fair. Fair is processed to predicted relativistic shift.

 

 
During the experiment, a series of error resources were introduced, which fall into two categories: the errors generated in the data acquisition process and Data processing progress.

4.Error of data acquisition

1) Time standard

a)Difference of the time standards    During the experiment, two sets of time standards were applied: one on the payload and the other in the ground base. The one on the payload generates the downlink signal and the comparison group. The one in the ground base is used in the first comparison with the downlink signal to generate the signal f1 – f0.    So the error caused by the difference between the two time standards were only introduced in the first order comparison.
During the pre-launch testing, the frequency of two time standards were adjusted. One of the zero beat was shown in figure .    b ) During the light time of downlink signal          i) Change in frequency of signal    When the EM wave travels in the troposphere(which covers the space within an altitude of 12 kilometres while the apogee of the rocket is within 2 kilometres), the velocity of light, c, is not a constant but a vector. It varies with graphic parameters.    Where n is the refractive index of the troposphere is given by equation [54]

Where T is the absolute temperature, p is the absolute pressure, ew is the absolute humidity(in terms of hPa)
Equation [53] is substituted into each sample, after this correction,  , is 5%.

ii) Path variation

The propagation vector in equation [1],  is assumed to be a straight line. However, due to the nonuniform variation of air particles, refraction take place. So the real propagation path of the downlink signal is a curve rater than a straight line. The actual path is given below.

Where rm is the displacement of the rocket from the centre of the earth, which is given
by  . Where  is the transformation matrix that map the navigation frame to the earth frame.

Where L and  are the longitude and latitude of the launch site.

After correction, the error in path determination can be ignored.

5.Error of frequency estimation

The received signal was discretized and the frequency of it was calculated though algorithms by digital circuits. During this process, errors are introduced. In this section, the error introduced by the frequency estimation algorithm will be discussed.
Assume the actual value of four parameters of the sample signal is ω0, A0, B0  and C0, the sampled value of the signal at time tk is y(k), the RSS of them is given by

△ω = ω – ω0
△A = A – A0
△B = B – B0
△C = C – C0
[60]

The difference in frequency between the estimation value and actual value is r(k) at time tk. The value of r(k) is given by

Substitute the equation [64] to the equation [63], equation [65] is obtained. This equation gives the error of this algorithm.

In this experiment, the sample rate was 120 million samples/s. The frequency was calculated per 10 second. The SNR of the signal was 50dB+. As a result, the accuracy of this algorithm reached a level of 1×10-12 Hz.

6.Conclusion

An experimental verification of the gravitational blueshift has been successfully achieved as a test of the general relativity conception.

作者简介:李雨翀,火箭爱好者,高中学生。初中时开始接触火箭制作,开发有RF系列火箭,已发射二十余次,是国内较早采用惯性导航技术的业余火箭之一。

利用软件设计液体燃料火箭发动机

经过一段时间对文献的研究,我们整理出一些列公式,然后根据公式利用VB编写了一个小程序,该程序适用于液体发动机的理论设计。再次感谢拔刀斋版主提供的Guipep热力学计算权威软件。

下面请跟着我一步一步来通过软件计算设计一个液体发动机要件的基本尺寸结构。
STEP1——设计要求:设计一个以90%质量浓度的过氧化氢为氧化剂,甲醇+乙醇胺+氯化铜(可溶性催化剂)为燃料的液体发动机。发动机燃烧室压力2MPa,设计推力20Kg(约200N)。根据氧平衡和点火性能,得出燃料比例为75:12:12:1(本文不作详细说明)。
STEP2——下载以下两个软件(Guipep以及液机尺寸设计均为绿色免安装软件,前者对visita系统兼容不好)
GUIPEP.rar (295 K)

液机尺寸设计.rar (6 K)
注:由于本站的下载功能尚未完成,下载需注册科创论坛,诚致歉意。
STEP3——找到Guipep.exe文件,双击运行之(该软件的dos计算内核为NASA开发,计算结果可信度较高,尤其适用于燃烧产物固相含量很少的液体发动机),出现该界面,如下图。


STEP4——在Desciption下选择各种原料对应的英文名称,并在右边的Weight内输入质量,单位g,质量设置为0表示取消该原料选择。在右边Title框中输入这种燃料的命名(自定),在Operating Conditions栏目下输入燃料常温(单位K)、燃烧室设计压强(单位PSI)、当地大气压强(单位PSI)。我处气温20℃(293K)、燃烧室压强2MPa(2×145=290PSI)、当地大气压为0.85atm(0.85×14.7PSI=12.5PSI),设置好的参数如下图。


STEP5——在工具栏中找到Run—>Singel Run单击之或者按下组合键Ctrl+R,此时弹出如下dos视窗。


STEP6——在dos视窗中按回车键执行,打印出一个文本文档,里面的内容即是计算结果。如下图


STEP7——找到“******CHAMBER RESULTS FOLLOW    ******”下的燃烧室计算结果,以及“****PERFORMANCE: FROZEN ON FIRST LINE, SHIFTING ON SECOND LINE****”下的冻结流数据。如下图

这里,我们只对以下四个计算结果感兴趣。分别是:

燃烧温度T(K)=2401K;

产物气体比热比γ CP/CV=1.1952;

产物气体的平均分子量THE MOLECULAR WEIGHT OF THE MIXTURE=20.413;

以及冻结流理论比冲IMPULSE=222.6。

根据需要将这四个结果记录下来。
STEP8——找到“液机尺寸设计.exe”文件并运行之,弹出如下界面。(该软件系自行编写,没时间去修正bug,例如除数为0错误等,各位海涵)


STEP9——根据设计任务的要求,将各项数据输入到对应的框中。
这里的氧燃比定义为氧化剂总质量比上燃料总质量;
特征长度定于为燃烧室容积除以喉口面积,表征了燃料在燃烧室内驻留的时间也既是燃烧完全程度,一般根据反应物活性取值1.5-6;
径收敛比定义为燃烧室直径与喉口直径的比值,表征了燃烧室喷注面积的大小,一般取3-7。
填好后,点击“计算”按钮进行计算,得到的结果如下图。


根据计算结果:
1.可以根据氧化剂流量和供液压力确定需要的氧化剂喷嘴型号和数目
2.可以根据燃料流量和供液压力确定需要的燃料喷嘴型号和数目
3.可以根据喷管排气速度判定是否为超音速发动机,以及火箭是否能够超音速飞行
4.可以根据喉口直径和出口直径设计拉瓦尔喷管扩张段尺寸
5.可以根据扩张比和喉口直径,来计算不同扩张角下的出口直径
6.可以直接根据燃烧室内径和长度设计发动机内壁尺寸

适当说明一下:

关于特征长度,取决于采用的喷注结构、燃料的汽化热和反应的活性。例如气态氧和汽油,可以低于1.5m,而过氧化氢多一个分解的过程,故取值较高。

关于径收敛比,取决于喷注面大小,也就是喷注流达到完全混合时所需要的最小面积,喷嘴雾化角越大,收敛比越大。
关于拉瓦尔喷管,较优性能和加工便捷度的设计为收敛段半角60°,喉口长度等于0.7倍喉口直径,扩张半角15°,三个部分均为圆弧过渡,不允许留有台阶或较深的刀纹。
关于金属壁厚与耐压的设计,参见此贴http://bbs.kechuang.org/read-kc-tid-31380.html

至此,一个液体发动机的燃烧室关键尺寸就计算出来了,可以根据这个数据设计图样进行加工。关于喷注器、材料、冷却系统、加工等具体说明,请参见本人译著《液体火箭发动机设计与制作》,链接http://bbs.kechuang.org/read-kc-tid-37433.html

原著:ehco魏

关于万户号多级火箭的技术建议

关于万户号多级火箭的技术建议

“沉默”

最近一直在关注有关“万户号”设想和方案的讨论,经收集一些网友的建议后,本人做了一点初步的设计,进行简单罗列。关于一级点火方式和降落伞开伞方案,相信已经有了比较成熟的方案,这里不再赘述。
发动机壳体用硬铝材料,外径48,壁厚4,前后两头用螺纹,并加工密封面。
堵头和喷管为碳钢,与壳体螺纹连接,用截面直径为3.55的O形圈与壳体密封。
关于一二级分离,提出了三种方案设想。
一种是二级点火将一级吹离,一级堵头和连接段一体化设计一级工作时推力由堵头传递至二级的喷管前端面。
二级喷管上有截面直径为1.8的O形圈,防止点火时气体从其他地方泄露。二级工作后,燃气压力将一级堵头与二级喷管分开,两者之间的行程可以设计在5mm左右,保证能够很快分离,且分离后一二级之间不会碰撞。

 另一种是利用弹射器将一二级弹开,弹射器装在一级堵头前面,一级堵头与二级之间有一个过渡舱,可以用铝合金,以减轻重量。二级喷管后面粘接一个2mm厚度的铝片,一级工作后通过延时方式给弹射器点火,推动弹射杆到铝片端面,并将二级推出密封面,之后两级之间分离。弹射器的行程可以设计在8mm左右,弹射杆与弹射器容腔之间有泄压孔,弹射杆运动到指定位置后,多余燃气从容腔上的泄压孔释放,弹射器、过渡舱与一级一起抛掉。二级点火后,喷管上的铝片被燃气吹掉。

还有一种是第二种方案的衍生,基本结构与第二种相同,弹射器改为弹簧结构,弹簧达到指定的压缩量后,将两端由不锈钢丝拉紧,钢丝的中间位置安装一个药包,工作时将药包点着,钢丝断开,弹簧向外弹出,推动二级喷管移动。

玉兔2A型火箭首次发射点火不成功原因分析

玉兔2号火箭首次实效发射试验中,点火不成功,火箭没有离开发射架。

现象:点火后发动机立即发火,发出一声喘气声,并喷出少量火焰,持续约2秒,随后熄火。

点火方式:等离子弧点火。

燃料:高氯酸铵—环氧树脂(RAP),含增塑剂。发动机为高强度合金钢材料。

试验前对该发动机使用的燃料抽样进行了开放点火试验,没有发现任何异常,燃烧充分、均匀,产气量大,无烟。

对故障发动机进行了分析,X光片(图1)显示,只有最靠近喷管的药柱被点燃,烧掉一个圆锥状药层后熄火。

(图1,工作后的发动机,高感度X光片,2mA,100KV,5min)

燃料质量虽然有一些瑕疵(图2),但不论是图1显示的瑕疵,还是图2显示的瑕疵,都不足以引起点火不成功。

(图2,第二行为工作后的药柱,2mA,65KV,5s,实际上除靠近喷管处的一节之外,其余药柱均未点燃)

燃料表面镜检,应该说是基本均匀的(图3),虽然有一些数十微米级别的大粒AP晶体碎片(图4),但也不足以引起点火不成功。

(图3,100倍镜检)

(图4,400倍镜检)

可以看到,药柱有少量瑕疵,按照以往经验,瑕疵可能导致燃料爆轰或发动机内压超过耐压而爆炸,但这些瑕疵不会引起点火不成功。

有一节药柱有较大缺陷(图5),但是点火并呈现锥形的药柱并不是这一节。其它药柱未发现类似缺陷。
(图5,药柱端面)

为了彻底找到原因,进行了地面试车试验。当采用0.5克左右的强力点火药粉制成点火头进行爆燃式点火时,点火药粉所在的燃料段正常燃烧,推力达到设计值,马赫环出现(图6)。其余燃料段缓慢燃烧,冒出亮黄色火焰,无推力(图7)。

(图6,点火后约0.5秒)

(图7,点火后约8秒)

结论:综合各方面因素,排除药柱瑕疵、发动机瑕疵、点火装置故障导致点火不成功的可能。该次试验点火不成功,是因为点火方式与燃料形式不匹配导致的。

作者:echo,虎哥,SS,焓熵,咸鱼,猴子,拉风,神话,风云等

RAP燃料是一种性能非常出色的燃料,但是工艺难度非常高。在图2最下面一排左起第三个药柱上,内孔上部阴影区域表示捏合不均,而捏合与固化是矛盾的,固化与真空排气也有矛盾。因此还需在工艺流程方面做进一步的完善。玉兔2号的航电部分在试验中工作正常。(撰文:刘虎)

讨论:http://bbs.kechuang.org/read-kc-tid-35590-fid-156.html

火箭燃料专用真空恒温浇铸仓的制作

文/ehco@科创论坛

论坛这个帖子的讨论在 http://bbs.kechuang.org/read-kc-tid-33071.html 进行,欢迎关注

我在试验AP基复合燃料的过程中,因为药柱铸造工艺不过关,曾导致多次发动机爆炸悲剧,在星空厂长指导下,慢慢研究起了真空浇铸技术,经过多次试验获得良好效果。在此将一些列工艺相关的问题作为系列和大家交流。

本文就燃料浇铸过程中必须用到的真空恒温仓和大家分享一下经验。
很多基于浇铸工艺的火箭燃料,都必须在真空恒温环境下进行浇铸。这样能最大程度地减少药柱中的气泡瑕疵,保证了良好的内弹道特性。

用合适尺寸的无缝钢管和钢板通过车床加工和焊接等工艺,制作两个能相互嵌套的罐子,内罐壁厚为 继续阅读“火箭燃料专用真空恒温浇铸仓的制作”

万户0型二级火箭初步方案

在科创航天计划中,万户系列火箭是多级固体推进剂小型火箭,最高可以配置到3级。开发万户系列火箭的主要目的是攻克多级探空火箭的关键技术。

万户0型火箭是万户系列的第一个实验型,主要用于测试级间分离技术、刹车技术、伞降技术。

万户0型火箭计划采用2级固体发动机。第一级粗短,第二级细长。第一级计划采用S.P燃料,工作时间小于0.5秒,最大加速度50g。第一级 继续阅读“万户0型二级火箭初步方案”