CN111516196B - 一种复合材料固体火箭发动机绝热层结构及其制备方法 - Google Patents
一种复合材料固体火箭发动机绝热层结构及其制备方法 Download PDFInfo
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Abstract
本发明提供一种复合材料固体火箭发动机绝热层结构及其制备方法,绝热层结构包括:绝热结构底层;绝热结构盖层,贴合设置于绝热结构底层的内侧;抗烧蚀结构层,设于绝热结构底层的喷管连接端,并贴合设置于绝热结构底层的外侧;金属接头,设于抗烧蚀结构层的外侧。制备方法包括:分别预制抗烧蚀结构层和绝热结构盖层;将金属接头和预制的抗烧蚀结构层组装后,在抗烧蚀结构层上设置胶片,整体成型预制绝热结构底层;将预制的绝热结构盖层设置在预制的绝热结构底层上,整体模压成型。该复合材料固体火箭发动机绝热层结构及其制备方法生产效率高,可保证绝热结构层各部分厚度可控,人工脱粘层深度准确,产品质量高。
Description
技术领域
本发明属于火箭发动机绝热结构制造技术领域,具体涉及一种复合材料固体火箭发动机绝热层结构及其制备方法。
背景技术
固体火箭发动机由燃烧室、推进剂、点火器和喷管等部件组成,其中,燃烧室是用于贮存推进剂、提供燃烧的空间,在火箭发动机工作过程中,燃烧室需要承受3000度以上的高温、高压气流冲刷,为避免燃烧室壳体受损,需要在燃烧室壳体内表面设置绝热层结构。
目前,国内制造固体火箭发动机内绝热层的方法主要有手工贴片成型和模压成型两种工艺。中国专利文献CN102632683B公开了一种纤维缠绕发动机壳体绝热层的手工贴片制作方法,通过手工贴片的方法制作发动机壳体绝热层,该方法可以保证纤维缠绕发动机壳体绝热层的手工贴片厚度尺寸,但是,手工贴片生产方式效率低和易出现鼓包、凹坑、脱粘等缺陷,极易受到热气流侵蚀,从而导致热防护失效。为此,整体模压成型技术得到开发和利用,其优点是可使手工贴片接缝至燃烧室筒体段,但是现代作战环境对导弹各项性能要求越来越高,导致绝热层的结构呈现为复杂化趋势,成型难度日益增加,模压成型工艺受制于固体火箭发动机的尺寸和绝热层的结构,传统的整体模压成型工艺难以满足发展的需要,无法保证绝热结构层的厚度、外形尺寸及内部质量。
发明内容
本发明解决的技术问题是提供一种复合材料固体火箭发动机绝热层结构及其制备方法,生产效率高,可保证绝热结构层各部分厚度可控,人工脱粘层深度准确,产品质量高。
为了解决上述问题,本发明提供一种复合材料固体火箭发动机绝热层结构,包括:
绝热结构底层,适于与发动机壳体封头内壁贴合设置;
绝热结构盖层,贴合设置于绝热结构底层的内侧;
抗烧蚀结构层,设于绝热结构底层的喷管连接端,并贴合设置于绝热结构底层的外侧;
金属接头,设于抗烧蚀结构层的外侧。
需要说明的是,本文中所述内侧、外侧,均以与发动机壳体封头的轴线的相对位置定义,靠近轴线的一侧为内侧。
优选地,绝热结构底层包括一体设置的直边段和椭球段,椭球段向下延伸形成接口段;直边段适于与发动机壳体封头内壁贴合设置;抗烧蚀结构层的一端与接口段的端部贴合,另一端延伸至椭球段与直边段的连接处;金属接头的一端与抗烧蚀结构层的端部贴合,另一端延伸至椭球段与直边段的连接处。
优选地,抗烧蚀结构层包括依次贴合设置的抗烧蚀底层、抗烧蚀层和抗烧蚀层盖层。
优选地,还包括接头保护层,接头保护层贴合设置于金属接头的外侧。
优选地,绝热结构底层的材料可选三元乙丙橡胶、丁腈橡胶。
优选地,绝热结构盖层的材料可选三元乙丙橡胶、丁腈橡胶。
优选地,抗烧蚀结构层中抗烧蚀底层的材料可选三元乙丙橡胶、丁腈橡胶;抗烧蚀层的材料可选碳纤维针刺预制体/改性酚醛、碳纤维三维编织/丁腈改性酚醛;抗烧蚀层盖层的材料可选三元乙丙橡胶、丁腈橡胶。
本发明的另一目的是提供一种制备上述的复合材料固体火箭发动机绝热层结构的方法,包括以下步骤:
根据选定的绝热层结构的厚度和外形尺寸分别预制抗烧蚀结构层和绝热结构盖层;将金属接头和预制的抗烧蚀结构层组装后,在抗烧蚀结构层上设置胶片,整体成型预制绝热结构底层;将预制的绝热结构盖层设置在预制的绝热结构底层上,整体模压成型,得到复合材料固体火箭发动机绝热层结构。
本发明采用模压硫化的方法制备成型发动机绝热层结构,相比于手工贴片生产方式,可大大提高生产效率和产品的质量;采用将抗烧蚀结构层、绝热结构底层、绝热结构盖层分别先预制,然后再整体模压成型的方法,可保证变厚度绝热结构层中各结构层的厚度、外型尺寸可控、内部质量可控,同时,绝热结构底层预制成型,可准确控制绝热结构底层与绝热结构盖层之间的人工脱粘层的深度,显著提高产品质量。
优选地,具体包括以下步骤:
S1.通过模压硫化成型预制抗烧蚀结构层,硫化阶段控制在硫化诱导期,得到抗烧蚀结构层预制件;
S2.在绝热结构底层模具的阴模中依次设置金属接头和抗烧蚀结构层预制件,然后在抗烧蚀结构层预制件上铺贴胶片,通过模压硫化整体成型预制绝热结构底层,硫化阶段控制在预硫化阶段,得到绝热结构底层预制件;
S3.通过模压硫化成型预制绝热结构盖层,硫化阶段控制在预硫化阶段,得到绝热结构盖层预制件;
S4.在绝热结构底层预制件上设置绝热结构盖层预制件,通过整体模压硫化成型,硫化阶段控制在正硫化阶段,得到复合材料固体火箭发动机绝热层结构。
本发明通过对绝热结构底层、绝热结构盖层分别预制,并将硫化阶段控制在预硫化阶段,在二者之间设置人工脱粘层之后,再整体正硫化成型,可防止绝热结构底层和绝热结构盖层过硫化而使其机械性能降低。
优选地,步骤S1具体包括以下步骤:
S1a.利用胶片通过模压硫化预制抗烧蚀底层,硫化阶段控制在硫化诱导期,得到抗烧蚀底层预制件;
S1b.利用胶片通过模压硫化预制抗烧蚀盖层,硫化阶段控制在硫化诱导期,得到抗烧蚀盖层预制件;
S1c.在抗烧蚀底层预制件、抗烧蚀盖层预制件的粘接面上涂刷胶粘剂,然后于抗烧蚀结构层模具中依次铺放抗烧蚀底层预制件、抗烧蚀层和抗烧蚀盖层预制件,通过整体模压硫化成型,硫化阶段控制在硫化诱导期,得到抗烧蚀结构层预制件。
本发明先对抗烧蚀结构层中的抗烧蚀盖层、抗烧蚀底层分别进行预制,然后再与抗烧蚀层整体预成型,可避免抗烧蚀盖层、抗烧蚀底层的生胶料流动塑形过程中裹气和胶黏剂的损失,显著提高抗烧蚀结构层与绝热结构底层、之间的界面强度,同时也可保证绝热结构底层与金属接头之间的界面强度。
优选地,还包括:
在步骤S2之前通过模压硫化成型预制接头保护层,硫化阶段控制在硫化诱导期,得到接头保护层预制件;步骤S2中,将接头保护层预制件与金属接头组装后,整体设置在绝热结构底层模具的阴模中。
优选地,步骤S4具体包括:
根据人工脱粘层的深度在绝热结构底层预制件上粘贴脱模胶带,形成人工脱粘层,将绝热结构盖层预制件与绝热结构底层预制件粘接,通过整体模压硫化成型,硫化阶段控制在正硫化阶段,得到复合材料固体火箭发动机绝热层结构。
优选地,步骤S4还包括:
在绝热结构盖层预制件与绝热结构底层预制件的粘接位置处设置生胶片。
优选地,生胶片为三元乙丙橡胶片。
优选地,步骤S4中,整体模压成型的方法为模具整体成型或气囊整体成型。
进一步优选地,步骤S4中,整体模压成型的方法为气囊整体成型。采用气囊整体成型,不仅解决了由于绝热结构底层、绝热结构盖层脱开深度终点置于模压方向垂直处,模具无法加压的问题,而且整体成型只需简单工装,节省成本效果明显。
本发明与现有技术相比,具有以下有益效果:
1.本发明的复合材料固体火箭发动机绝热层结构及其制备方法,采用模压硫化的方法制备成型发动机绝热层结构,相比于手工贴片生产方式,可大大提高生产效率和产品的质量;采用将抗烧蚀结构层、绝热结构底层、绝热结构盖层分别先预制,然后再整体模压成型的方法,可保证变厚度绝热结构层中各结构层的厚度、尺寸可控,并可准确控制绝热结构底层与绝热结构盖层之间的人工脱粘层的深度,显著提高产品质量;
2.本发明的复合材料固体火箭发动机绝热层结构及其制备方法,通过对绝热结构底层、绝热结构盖层分别预制,并将硫化阶段控制在预硫化阶段,在二者之间设置人工脱粘层之后,再整体正硫化成型,可防止绝热结构底层和绝热结构盖层过硫化而使其机械性能降低;
3.本发明的复合材料固体火箭发动机绝热层结构及其制备方法,先对抗烧蚀结构层中的抗烧蚀盖层、抗烧蚀底层分别进行预制,然后再与抗烧蚀层整体预成型,可避免抗烧蚀盖层、抗烧蚀底层的生胶料流动塑形过程中裹气和胶黏剂的损失,显著提高抗烧蚀结构层与绝热结构底层、之间的界面强度,同时也可保证绝热结构底层与金属接头之间的界面强度;
4.本发明的复合材料固体火箭发动机绝热层结构及其制备方法,采用气囊整体成型,不仅解决了由于绝热结构底层、绝热结构盖层脱开深度终点置于模压方向垂直处,模具无法加压的问题,而且整体成型只需简单工装,节省成本效果明显。
附图说明
图1是本发明实施例所述的复合材料固体火箭发动机绝热层结构的结构示意图;
图2是本发明实施例所述的复合材料固体火箭发动机绝热层结构的分解示意图;
图3是本发明实施例所述的复合材料固体火箭发动机绝热层结构中抗烧蚀结构层的结构示意图。
其中:1-绝热结构底层;11-直边段;12-椭球段;13-接口段;2-绝热结构盖层;3-抗烧蚀结构层;31-抗烧蚀底层;32-抗烧蚀层;33-抗烧蚀层盖层;4-金属接头;5-接头保护层。
具体实施方式
下面将结合本发明的实施例,对本发明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。
实施例一
如图1、图2、图3所示,本实施例所述的一种复合材料固体火箭发动机绝热层结构,包括:
绝热结构底层1,适于与发动机壳体封头内壁贴合设置,包括一体设置的直边段11和椭球段12,椭球段12向下延伸形成接口段13;直边段11适于与发动机壳体封头内壁贴合设置,绝热结构底层所使用的材料为三元乙丙橡胶;
绝热结构盖层2,贴合设置于绝热结构底层的内侧,绝热结构盖层所使用的材料为三元乙丙橡胶;
抗烧蚀结构层3,设于绝热结构底层1的喷管连接端,并贴合设置于绝热结构底层1的外侧,抗烧蚀结构层3的一端与接口段13的端部贴合,另一端延伸至椭球段12与直边段11的连接处,抗烧蚀结构层3包括依次贴合设置的抗烧蚀底层31、抗烧蚀层32和抗烧蚀层盖层33,抗烧蚀底层31使用的材料为三元乙丙橡胶;抗烧蚀层32使用的材料为碳纤维针刺预制体/改性酚醛;抗烧蚀层盖层33使用的材料为三元乙丙橡胶;
金属接头4,设于抗烧蚀结构层3的外侧,一端与抗烧蚀结构层3的端部贴合,另一端延伸至椭球段12与直边段11的连接处;
接头保护层5,贴合设置于金属接头4的外侧,接头保护层5使用的材料为三元乙丙橡胶。
本实施例的制备上述的复合材料固体火箭发动机绝热层结构的方法,包括以下步骤:
S0.使用压延机将各层结构的生胶料碾压成合理厚度胶片,根据绝热结构层三维模型,利用计算机软件对其展开,根据碾片机设备幅宽确定尺寸,将胶片裁剪至合理外形尺寸;
S1a.将抗烧蚀底层胶片铺设在抗烧蚀底层模具中,将模具升温至80-90℃保温30min,使硫化阶段控制在硫化诱导期,通过模压硫化预制抗烧蚀底层,得到抗烧蚀底层预制件;
S1b.将抗烧蚀盖层胶片铺设在抗烧蚀盖层模具中,将模具升温至80-90℃保温30min,使硫化阶段控制在硫化诱导期,通过模压硫化预制抗烧蚀盖层,得到抗烧蚀盖层预制件;
S1c.在抗烧蚀底层预制件、抗烧蚀盖层预制件的粘接面打磨见新,清理干净,并涂刷胶粘剂,然后于抗烧蚀结构层模具中依次铺放抗烧蚀底层预制件、抗烧蚀层和抗烧蚀盖层预制件,将模具升温至85-95℃保温30min,使硫化阶段控制在硫化诱导期,通过整体模压硫化成型,得到抗烧蚀结构层预制件;
S2a.将接头保护层胶片铺设在接头保护层模具中,将模具升温至80-90℃保温30min,使硫化阶段控制在硫化诱导期,通过模压硫化预制接头保护层,得到接头保护层预制件;
S2b.将绝热结构底层模具堵头装于绝热结构底层模具阴模中,使用螺栓连接,在接头保护层预制件上涂刷胶粘剂,并铺放于金属接头上,使用抱环将其固定,组合件整体放置在绝热结构底层模具阴模中,在金属接头上铺贴抗烧蚀结构层预制件,安装绝热结构底层模具堵头配合模具,根据设计厚度在阴模中进行绝热结构底层胶片铺层,并用胶黏剂进行固定,铺层完成后将绝热结构底层阳模与阴模合模,置于成型设备上135-145℃保温90min,使硫化阶段控制在预硫化阶段,通过模压硫化整体成型,得到绝热结构底层预制件,然后脱去绝热结构底层阳模和堵头配合模具;
S3.将绝热结构盖层胶片铺设在绝热结构盖层模具中,将模具升温至135-145℃保温60min,使硫化阶段控制在预硫化阶段,通过模压硫化成型,得到绝热结构盖层预制件;
S4.根据设定的人工脱粘层的深度,在绝热结构底层预制件上脱开深度内粘接脱模胶带,形成人工脱粘层,将绝热结构盖层预制件和绝热结构底层预制件粘接位置打磨清理,并涂刷胶粘剂,在绝热结构底层阴模中,绝热结构底层预制件上放置绝热结构盖层预制件,并在二者粘接位置处设置生三元乙丙橡胶片,装入整体成型模具气囊下封工装,置入气囊,将其气嘴伸入整体成型模具阳模内,将模具整体置于成型设备中140-150℃保温90min硫化成型,气囊充气增加压力,使硫化阶段控制在正硫化阶段,然后自然降温至60℃以下,脱模,得到复合材料固体火箭发动机绝热层结构。
经过多次采用手工贴片制备和按照上述方法制备复合材料固体火箭发动机绝热层结构,总结实验数据得出,采用手工贴片制备得到的绝热层结构,盖层厚度范围为1.8mm~3.2mm,底层厚度为6.7~10.4mm,厚度控制精度较差,而采用本实施例的方法可将盖层厚度控制在(2±0.3)mm,底层厚度控制在(8±0.5)mm,人工脱粘层深度也可实现精准控制,满足(251±1)mm的设计要求。
对采用手工贴片的方法制备和采用本实施例的方法制备得到的复合材料固体火箭发动机绝热层结构的抗烧蚀层、金属接头界面粘接情况采用超声方法进行检测,内部质量合格率由手工贴片方法的63%提升至98%。
对采用手工贴片的方法制备和采用本实施例的方法制备得到的复合材料固体火箭发动机绝热层结构的金属接头、抗烧蚀层、绝热层的界面强度进行测定。代表其界面强度的随炉试样,抗烧蚀层采用(金属/绝热材料/抗烧蚀层/绝热材料/金属)试样,金属接头界面采用(金属/绝热材料/金属)试样,按照QJ2038.1A和QJ2038.2标准测试。金属/绝热材料/抗烧蚀层/绝热材料/金属试样,扯离强度从1.83MPa提高至2.51MPa,剪切强度从1.52MPa提高至1.95MPa;金属/绝热材料/金属试样,扯离强度从2.18MPa提升至3.21MPa,剪切强度从1.98MPa,提升至2.64MPa。
显然,上述实施例仅仅是为清楚地说明所作的举例,而并非对实施方式的限定。对于所属领域的普通技术人员来说,在上述说明的基础上还可以做出其它不同形式的变化或变动。这里无需也无法对所有的实施方式予以穷举。而由此所引伸出的显而易见的变化或变动仍处于本发明创造的保护范围之中。
Claims (6)
1.一种复合材料固体火箭发动机绝热层结构,其特征在于,包括:
绝热结构底层,适于与发动机壳体封头内壁贴合设置;
绝热结构盖层,贴合设置于所述绝热结构底层的内侧;
抗烧蚀结构层,设于所述绝热结构底层的喷管连接端,并贴合设置于所述绝热结构底层的外侧,所述抗烧蚀结构层包括依次贴合设置的抗烧蚀底层、抗烧蚀层和抗烧蚀盖层;
金属接头,设于所述抗烧蚀结构层的外侧;
所述复合材料固体火箭发动机绝热层结构的制备方法包括以下步骤:
S1a.利用胶片通过模压硫化预制抗烧蚀底层,硫化阶段温度为80-90℃,保温30min,控制在硫化诱导期,得到抗烧蚀底层预制件;
S1b.利用胶片通过模压硫化预制抗烧蚀盖层,硫化阶段温度为80-90℃,保温30min,控制在硫化诱导期,得到抗烧蚀盖层预制件;
S1c.在所述抗烧蚀底层预制件、所述抗烧蚀盖层预制件的粘接面上涂刷胶粘剂,然后于抗烧蚀结构层模具中依次铺放所述抗烧蚀底层预制件、抗烧蚀层和所述抗烧蚀盖层预制件,通过整体模压硫化成型,硫化阶段温度为85-95℃,保温30min,控制在硫化诱导期,得到所述抗烧蚀结构层预制件;
S2.在绝热结构底层模具的阴模中依次设置所述金属接头和所述抗烧蚀结构层预制件,然后在所述抗烧蚀结构层预制件上铺贴胶片,通过模压硫化整体成型预制所述绝热结构底层,硫化阶段温度为135-145℃,保温90min,控制在预硫化阶段,得到绝热结构底层预制件;
S3.通过模压硫化成型预制所述绝热结构盖层,硫化阶段温度为135-145℃,保温60min,控制在预硫化阶段,得到绝热结构盖层预制件;
S4.在所述绝热结构底层预制件上设置所述绝热结构盖层预制件,通过整体模压硫化成型,硫化阶段温度为140-150℃,保温90min,控制在正硫化阶段,得到所述复合材料固体火箭发动机绝热层结构;整体模压成型的方法为气囊整体成型。
2.根据权利要求1所述的复合材料固体火箭发动机绝热层结构,其特征在于:
所述绝热结构底层包括一体设置的直边段和椭球段,所述椭球段向下延伸形成接口段;所述直边段适于与所述发动机壳体封头内壁贴合设置;
所述抗烧蚀结构层的一端与所述接口段的端部贴合,另一端延伸至所述椭球段与所述直边段的连接处;
所述金属接头的一端与所述抗烧蚀结构层的端部贴合,另一端延伸至所述椭球段与所述直边段的连接处。
3.根据权利要求1所述的复合材料固体火箭发动机绝热层结构,其特征在于:
还包括接头保护层,所述接头保护层贴合设置于所述金属接头的外侧。
4.根据权利要求1所述的复合材料固体火箭发动机绝热层结构,其特征在于,制备方法还包括:
在步骤S2之前通过模压硫化成型预制接头保护层,硫化阶段控制在硫化诱导期,得到接头保护层预制件;步骤S2中,将所述接头保护层预制件与所述金属接头组装后,整体设置在所述绝热结构底层模具的阴模中。
5.根据权利要求1所述的复合材料固体火箭发动机绝热层结构,其特征在于,步骤S4具体包括:
根据人工脱粘层的深度在所述绝热结构底层预制件上粘贴脱模胶带,形成人工脱粘层,将所述绝热结构盖层预制件与所述绝热结构底层预制件粘接,通过整体模压硫化成型,硫化阶段控制在正硫化阶段,得到所述复合材料固体火箭发动机绝热层结构。
6.根据权利要求1所述的复合材料固体火箭发动机绝热层结构,其特征在于,步骤S4还包括:
在所述绝热结构盖层预制件与所述绝热结构底层预制件的粘接位置处设置生胶片。
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