CN112192862A - 一种复合材料火箭发动机壳体制备方法 - Google Patents
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Abstract
本发明公开了一种复合材料火箭发动机壳体制备方法,包括裙,所述裙的一端固定连接有承力结构层,所述承力结构层的外侧面固定安装有绝热层,所述承力结构层及绝热层的两端端面中间位置均贯穿设置有封头。本发明能够采用气囊加特制模具用气胀法成型内层,突破现有成型模具方案对产品的尺寸限制,实现各种极孔尺寸,长度尺寸的复合材料火箭发动机壳体制备,现有模具方案产生装配公差或机械加工带来的公差最终会反馈到产品上,采用本制备方法成型过程,能极大的避免公差累积,减小制备过程对产品成型质量的负面影响,以产品本身作为缠绕芯模,取消了现有方案芯模的装配和拆卸步骤,精简产品制备工艺流程,缩短生产周期。
Description
技术领域
本发明涉及航天设备制造领域,尤其是涉及一种复合材料火箭发动机壳体结构及其制备方法。
背景技术
随着航天产业的高速发展,对火箭的运载能力需求越来越高,降低重量、提高容积成为火箭发动机壳体的主要方向,高性能碳纤维复合材料作为一种新型火箭发动机壳体制造材料,在等强度的条件下,重量是金属材料的三分之一到二分之一,不仅可以大幅度降低发射成本,而且可以有效提高发射行程,具有很大的经济和战略意义。
复合材料火箭发动机壳体制造的核心是成型模具的设计和制备,目前中小型筒体缠绕成型模具通常采用两种方案:1、金属芯模结构,采用壳板、芯轴和支撑件的形式装配组成模具,产品成型后,操作者从极孔处进入产品腔体内部进行拆卸,零件从产品极孔处运送出来,该方案一般运用于大尺寸壳体;2、可溶性芯模结构,采用颗粒状材料通过一定手段聚合形成模具,产品成型后,溶解模具恢复颗粒状,从极孔中取出,小尺寸壳体多采用此模具形式。
目前的两种方案都有一定的局限性,若产品极孔尺寸小,即无法提供操作者进出通道,则方案1不适用;若产品长度尺寸大,即模具要求较大的刚度,则方案2不适用,现有的方案使得小极孔长尺寸的火箭发动机壳体结构设计和制造存在极大地困难。
为此,提出一种复合材料火箭发动机壳体制备方法。
发明内容
本发明的目的在于提供一种复合材料火箭发动机壳体制备方法,采用气囊加特制模具用气胀法成型内层,突破现有成型模具方案对产品的尺寸限制,实现各种极孔尺寸,长度尺寸的复合材料火箭发动机壳体制备,以解决上述背景技术中提出的问题。
为实现上述目的,本发明提供如下技术方案:一种复合材料火箭发动机壳体结构,包括裙,所述裙的一端固定连接有承力结构层,所述承力结构层的外侧面固定安装有绝热层,所述承力结构层及绝热层的两端端面中间位置均贯穿设置有封头。
一种复合材料火箭发动机壳体制备方法,该制备方法包括如下步骤:
S1:制备裙:在钢制开放式阳模的模具上涂刷XTEND 19RSS脱模剂,手工铺设一定层数的T700碳纤维/环氧树脂150℃固化环氧预浸布料,铺层方向为±45°和0°/90°交错铺贴,第一层铺放完毕抽真空预压实,之后每4-5层抽真空,保证布层平整紧密的缠在模具上,在铺至理论厚度后增加两层铺层作为机加余量,之后使用热压罐加压抽真空固化;
S2:封头采用阴阳组合模具模压的方式制造,以脱开层作为工艺制造分离面,盖层和底层分别模压硫化,检测合格后再合在一起与金属接头共同硫化,完成封头零件,最终硫化时,已硫化的盖层和底层胶接面铺贴一层1mm厚的未硫化的丁腈料片,采用气囊加压硫化,固化后整形交检,金属接头为外购件,采用铝合金锻件,数控机加制造,采用阳极化处理,处理后密封存放并在规定时间内完成胶接,阳极化处理后最好在8小时内涂上底涂剂,涂刷底涂剂后应在24小时内进行胶接;
S3:在模具下模中依次铺贴一定层数的T700碳纤维/环氧树脂150℃固化环氧预浸布料、丁腈绝热材料,在模具两端放置预先制备的前后封头,接头定位和使用模具本身的台阶,在封头和模具中段的空腔内放置气囊,充气保压在500Pa,使气囊膨胀出理论形状,然后在气囊上表面继续依次铺贴一定层数的丁腈绝热材料、T700碳纤维/环氧树脂150℃固化环氧预浸布料,形成封闭腔体包裹住气囊和前后封头,再盖好上模,合模后采用标准紧固件锁死上下模具,然后放入烘箱固化;
S4:取出S3制备好的产品内胆,从封头极孔处抽去气囊,在内胆两端头外侧壁涂抹P700-1M缠绕树脂后套上S1制备好的裙,粘接后将内胆连同裙通过转接件固定在缠绕机上,等待缠绕,缠绕层材料为T700级碳纤维及P700-1M缠绕树脂,使用20束普通碳丝加单束45N张力,缠绕过程中需人工刮除纤维上的多余胶液,缠绕过程需采用专用工装进行直径测量,测量次数、位置及尺寸要求按缠绕过程测量图,缠绕完成后,按预固化,最终固化曲线完成产品结构的完全固化,固化后测量产品各尺寸是否满足要求,可打磨或补喷对内腔形状和尺寸进行微调,可机械加工对端面尺寸进行微调。
优选的,所述S1中固化制度为160℃保温4h,压力为1MPa,加压的目的是克服固化过程中产生的挥发份、去除材料中产生的孔隙和增加物料的流动性。
优选的,所述S1中铺层可对接,可打剪口,在对接缝>5mm时可用窄布条填充。
优选的,所述S2中封头分为前封头及后封头,前封头使用的绝热材料为丁腈绝热橡胶,后封头使用丁腈、5-III绝热材料。
优选的,所述S3中气囊采用硅橡胶制作,能保证1.5MPa压力下的气密性,气囊形状大小与产品内腔形状大小相同。
优选的,所述裙与承力结构层共胶接,承力结构层和绝热层使用缠绕树脂胶接,绝热层和封头采用底涂胶配合胶粘剂共同固化。
与现有技术相比,本发明的有益效果是:
1、预成型产品内胆,以产品本身作为缠绕芯模,取消了现有方案芯模的装配和拆卸步骤,精简产品制备工艺流程,缩短生产周期;
2、采用气囊加特制模具用气胀法成型内层,突破现有成型模具方案对产品的尺寸限制,实现各种极孔尺寸,长度尺寸的复合材料火箭发动机壳体制备;
3、现有模具方案产生装配公差或机械加工带来的公差最终会反馈到产品上,采用本制备方法成型过程,能极大的避免公差累积,减小制备过程对产品成型质量的负面影响;
4、本申请中的工艺流程门槛低、人员掌握熟练、培训容易、无需采购调试设备,所需工装模具大幅减少,极大地缩短生产周期,降低生产成本。
附图说明
为了更清楚地说明本发明具体实施方式或现有技术中的技术方案,下面将对具体实施方式或现有技术描述中所需要使用的附图作简单地介绍,显而易见地,下面描述中的附图是本发明的一些实施方式,对于本领域普通技术人员来讲,在不付出创造性劳动的前提下,还可以根据这些附图获得其他的附图。
图1为本发明的制备流程结构示意图;
图2为本发明的整体结构示意图;
图3为本发明的剖面结构示意图。
附图标记说明:
图中:1、裙;2、承力结构层;3、绝热层;4、封头。
具体实施方式
下面将结合本发明实施例中的附图,对本发明实施例中的技术方案进行清楚、完整地描述,显然,所描述的实施例仅仅是本发明一部分实施例,而不是全部的实施例。基于本发明中的实施例,本领域普通技术人员在没有做出创造性劳动前提下所获得的所有其他实施例,都属于本发明保护的范围。
在本发明的描述中,需要理解的是,术语"中心"、"纵向"、"横向"、"长度"、"宽度"、"厚度"、"上"、"下"、"前"、"后"、"左"、"右"、"坚直"、"水平"、"顶"、"底"、"内"、"外"、"顺时针"、"逆时针"等指示的方位或位置关系为基于附图所示的方位或位置关系,仅是为了便于描述本发明和简化描述,而不是指示或暗示所指的装置或元件必须具有特定的方位、以特定的方位构造和操作,因此不能理解为对本发明的限制。
此外,术语"第一"、"第二"仅用于描述目的,而不能理解为指示或暗示相对重要性或者隐含指明所指示的技术特征的数量。由此,限定有"第一"、"第二"的特征可以明示或者隐含地包括一个或者更多个所述特征。在本发明的描述中,"多个"的含义是两个或两个以上,除非另有明确具体的限定。此外,术语“安装”、“相连”、“连接”应做广义理解,例如,可以是固定连接,也可以是可拆卸连接,或一体地连接;可以是机械连接,也可以是电连接;可以是直接相连,也可以通过中间媒介间接相连,可以是两个元件内部的连通。对于本领域的普通技术人员而言,可以具体情况理解上述术语在本发明中的具体含义。
请参阅图1至图3,本发明提供一种复合材料火箭发动机壳体制备方法技术方案:
一种复合材料火箭发动机壳体结构,包括裙1,所述裙1的一端固定连接有承力结构层2,所述承力结构层2的外侧面固定安装有绝热层3,所述承力结构层2及绝热层3的两端端面中间位置均贯穿设置有封头4,所述裙1与承力结构层2共胶接,承力结构层2和绝热层3使用缠绕树脂胶接,绝热层3和封头4采用底涂胶配合胶粘剂共同固化;
如图1所示,一种复合材料火箭发动机壳体制备方法,该制备方法包括如下步骤:
S1:制备裙1:在钢制开放式阳模的模具上涂刷XTEND 19RSS脱模剂,手工铺设一定层数的T700碳纤维/环氧树脂150℃固化环氧预浸布料,铺层方向为±45°和0°/90°交错铺贴,第一层铺放完毕抽真空预压实,之后每4-5层抽真空,保证布层平整紧密的缠在模具上,在铺至理论厚度后增加两层铺层作为机加余量,之后使用热压罐加压抽真空固化,所述S1中固化制度为160℃保温4h,压力为1MPa,加压的目的是克服固化过程中产生的挥发份、去除材料中产生的孔隙和增加物料的流动性,所述S1中铺层可对接,可打剪口,在对接缝>5mm时可用窄布条填充;
S2:制备前后封头4:封头4采用阴阳组合模具模压的方式制造,以脱开层作为工艺制造分离面,盖层和底层分别模压硫化,检测合格后再合在一起与金属接头共同硫化,完成封头4零件,最终硫化时,已硫化的盖层和底层胶接面铺贴一层1mm厚的未硫化的丁腈料片,采用气囊加压硫化,固化后整形交检,金属接头为外购件,采用铝合金锻件,数控机加制造,采用阳极化处理,处理后密封存放并在规定时间内完成胶接,阳极化处理后最好在8小时内涂上底涂剂,涂刷底涂剂后应在24小时内进行胶接,所述S2中封头4分为前封头及后封头,前封头使用的绝热材料为丁腈绝热橡胶,后封头使用丁腈、5-III绝热材料;
S3:制备产品内胆:在模具下模中依次铺贴一定层数的T700碳纤维/环氧树脂150℃固化环氧预浸布料、丁腈绝热材料,在模具两端放置预先制备的前后封头4,接头定位和使用模具本身的台阶,在封头4和模具中段的空腔内放置气囊,充气保压在500Pa,使气囊膨胀出理论形状,然后在气囊上表面继续依次铺贴一定层数的丁腈绝热材料、T700碳纤维/环氧树脂150℃固化环氧预浸布料,形成封闭腔体包裹住气囊和前后封头4,再盖好上模,合模后采用标准紧固件锁死上下模具,然后放入烘箱固化,所述S3中气囊采用硅橡胶制作,能保证1.5MPa压力下的气密性,气囊形状大小与产品内腔形状大小相同;
S4:缠绕固化:取出S3制备好的产品内胆,从封头4极孔处抽去气囊,在内胆两端头外侧壁涂抹P700-1M缠绕树脂后套上S1制备好的裙1,粘接后将内胆连同裙1通过转接件固定在缠绕机上,等待缠绕,缠绕层材料为T700级碳纤维及P700-1M缠绕树脂,使用20束普通碳丝加单束45N张力,缠绕过程中需人工刮除纤维上的多余胶液,缠绕过程需采用专用工装进行直径测量,测量次数、位置及尺寸要求按缠绕过程测量图,缠绕完成后,按预固化,最终固化曲线完成产品结构的完全固化,固化后测量产品各尺寸是否满足要求,可打磨或补喷对内腔形状和尺寸进行微调,可机械加工对端面尺寸进行微调。
本发明采用气囊加特制模具用气胀法成型内层,突破现有成型模具方案对产品的尺寸限制,实现各种极孔尺寸,长度尺寸的复合材料火箭发动机壳体制备。
尽管已经示出和描述了本发明的实施例,对于本领域的普通技术人员而言,可以理解在不脱离本发明的原理和精神的情况下可以对这些实施例进行多种变化、修改、替换和变型,本发明的范围由所附权利要求及其等同物限定。
Claims (7)
1.一种复合材料火箭发动机壳体结构,包括裙(1),其特征在于:所述裙(1)的一端固定连接有承力结构层(2),所述承力结构层(2)的外侧面固定安装有绝热层(3),所述承力结构层(2)及绝热层(3)的两端端面中间位置均贯穿设置有封头(4)。
2.根据权利要求1所述的一种复合材料火箭发动机壳体制备方法,其特征在于:该制备方法包括如下步骤:
S1:制备裙(1):在钢制开放式阳模的模具上涂刷XTEND 19RSS脱模剂,手工铺设一定层数的T700碳纤维/环氧树脂150℃固化环氧预浸布料,铺层方向为±45°和0°/90°交错铺贴,第一层铺放完毕抽真空预压实,之后每4-5层抽真空,保证布层平整紧密的缠在模具上,在铺至理论厚度后增加两层铺层作为机加余量,之后使用热压罐加压抽真空固化;
S2:制备前后封头(4):封头(4)采用阴阳组合模具模压的方式制造,以脱开层作为工艺制造分离面,盖层和底层分别模压硫化,检测合格后再合在一起与金属接头共同硫化,完成封头(4)零件,最终硫化时,已硫化的盖层和底层胶接面铺贴一层1mm厚的未硫化的丁腈料片,采用气囊加压硫化,固化后整形交检,金属接头为外购件,采用铝合金锻件,数控机加制造,采用阳极化处理,处理后密封存放并在规定时间内完成胶接,阳极化处理后最好在8小时内涂上底涂剂,涂刷底涂剂后应在24小时内进行胶接;
S3:制备产品内胆:在模具下模中依次铺贴一定层数的T700碳纤维/环氧树脂150℃固化环氧预浸布料、丁腈绝热材料,在模具两端放置预先制备的前后封头(4),接头定位和使用模具本身的台阶,在封头(4)和模具中段的空腔内放置气囊,充气保压在500Pa,使气囊膨胀出理论形状,然后在气囊上表面继续依次铺贴一定层数的丁腈绝热材料、T700碳纤维/环氧树脂150℃固化环氧预浸布料,形成封闭腔体包裹住气囊和前后封头(4),再盖好上模,合模后采用标准紧固件锁死上下模具,然后放入烘箱固化;
S4:缠绕固化:取出S3制备好的产品内胆,从封头(4)极孔处抽去气囊,在内胆两端头外侧壁涂抹P700-1M缠绕树脂后套上S1制备好的裙(1),粘接后将内胆连同裙(1)通过转接件固定在缠绕机上,等待缠绕,缠绕层材料为T700级碳纤维及P700-1M缠绕树脂,使用20束普通碳丝加单束45N张力,缠绕过程中需人工刮除纤维上的多余胶液,缠绕过程需采用专用工装进行直径测量,测量次数、位置及尺寸要求按缠绕过程测量图,缠绕完成后,按预固化,最终固化曲线完成产品结构的完全固化,固化后测量产品各尺寸是否满足要求,可打磨或补喷对内腔形状和尺寸进行微调,可机械加工对端面尺寸进行微调。
3.根据权利要求2所述的一种复合材料火箭发动机壳体制备方法,其特征在于:所述S1中固化制度为160℃保温4h,压力为1MPa,加压的目的是克服固化过程中产生的挥发份、去除材料中产生的孔隙和增加物料的流动性。
4.根据权利要求2所述的一种复合材料火箭发动机壳体制备方法,其特征在于:所述S1中铺层可对接,可打剪口,在对接缝>5mm时可用窄布条填充。
5.根据权利要求2所述的一种复合材料火箭发动机壳体制备方法,其特征在于:所述S2中封头(4)分为前封头及后封头,前封头使用的绝热材料为丁腈绝热橡胶,后封头使用丁腈、5-III绝热材料。
6.根据权利要求2所述的一种复合材料火箭发动机壳体制备方法,其特征在于:所述S3中气囊采用硅橡胶制作,能保证1.5MPa压力下的气密性,气囊形状大小与产品内腔形状大小相同。
7.根据权利要求1所述的一种复合材料火箭发动机壳体制备方法,其特征在于:所述裙(1)与承力结构层(2)共胶接,承力结构层(2)和绝热层(3)使用缠绕树脂胶接,绝热层(3)和封头(4)采用底涂胶配合胶粘剂共同固化。
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