分段装药固体火箭发动机限燃层烧蚀实验方法

Experimental method for ablation of inhibitor in segment-propellant solid rocket motor

发展了一种能开展分段装药发动机限燃层烧蚀研究的模拟实验方法。研制了限燃层烧蚀实验装置,对实验装置及全尺寸分段装药固体火箭发动机的三维两相流内流场开展了对比数值模拟。结果证明,这种实验装置模拟的限燃层附近粒子低浓度及高速度环境与全尺寸分段装药发动机限燃层附近很接近,说明这种实验方法是可用于研究分段装药发动机限燃层烧蚀的。然后,利用此实验装置开展了限燃层烧蚀实验。实验结果表明,限燃层烧蚀后的马蹄形貌和粒子流的分布吻合;粒子侵蚀在限燃层烧蚀中占主要地位,且粒子浓度是影响烧蚀率的重要因素;烧蚀发动机限燃层烧蚀率与全尺寸发动机限燃层烧蚀率基本一致,分别为1.64 mm/s和1.52 mm/s,即此装置用于研究限燃层烧蚀切实可行;限燃层的烧蚀是从轴向到径向逐渐发展的,即轴向烧蚀率大于径向烧蚀率。

A simulation experimental method was developed to investigate the ablation characteristic of the inhibitor in segment- propellant solid rocket motor. A new inhibitor erosion facility was designed to do experiment investigation. Three-dimensional two- phase numerical simulations were carried out to compare the thermo environment of this test facility with the full-scale solid rocket motor around the inhibitor. The numerical results show that the facility can be used to simulate the environment of low particle con- centration and high velocity to investigate the inhibitor erosion. Then the experiment was carried out with this device, and the test results show that the inhibitor presents horseg hoof after erosion, which is coordinated with the distribution of particles ; particles ero- sion plays an important role of inhibitor ablation, and particles concentration is an important factor of ablation characteristic of inhibitor ; inhibitor erosion rate is similar with that of full-scale segment-propellant solid motor, which is 1.64 mm/s and 1.52 mm/s respectively;the erosion of inhibitor starts from the axial section and then develops to the radial section, namely the axial erosion rate is faster than radial erosion rate.