Multifunction Rocket System Development based on Advanced Hybrid Propulsion

Transcription

1 SpaceOps Conferences May 2016, Daejeon, Korea SpaceOps 2016 Conference / Multifunction Rocket System Development based on Advanced Hybrid Propulsion Yen-Sen Chen, Robert Cheng, Hao-Chi Chang, Luke Yang, Bill Wu National Space Organization, Hsinchu, Taiwan Alfred Lai, Jhe-Wei Lin, Shih-Sin Wei, Tzu-Hao Chou, Tsung-Lin Chen, Jong-Shinn Wu Dept. of Mech. Engr., National Chiao Tung University, Hsinchu, Taiwan and Ming-Tzu Ho Dept. of Engr. Science, National Cheng Kung University, Tainan, Taiwan Mainly due to its propellant non-mixing feature, hybrid rocket propulsion has been demonstrated to be more advantageous in operation safety as compared to its solid and liquid counterparts. The traditionally moderate Isp output of hybrid rockets has been enhanced to be close to the liquid rocket performance in recent years, particularly with the innovative designs employed in this research by using dual-vortical-flow (DVF) chambers. Based on this new approach and cost saving strategy, a multifunction rocket system is designed with the features of high performance hybrid combustion, trajectory following flight controls, enhanced science experiments, and an advanced payload recovery method. High fidelity numerical modeling design approach and hot-fire experiments are employed to assess the overall performance of the DVF hybrid rocket engines that has roll control capability embedded in the engine design. The present hybrid rocket engine designs consider propellant systems of N 2 O/HTPB, N 2 O/HDPE and H 2 O 2 /HDPE. Pressure-fed system is the baseline for delivering the oxidizer to the combustion chamber while pump-fed system is also considered as a design option, especially for the hydrogen peroxide system. Carbon fiber filament winding pressure tank is incorporated to contain the oxidizer. Pressurant is also employed for better thrust control. To enhance the overall performance and benefits of conducting flight experiments using hybrid rocket, three basic flight trajectory designs are proposed in this study, namely the traditional standard parabolic trajectory, a TASE (Trajectory Augmented Science Experiments) maneuver and a HOOK (Homing Oriented Operation Kernel) maneuver. The TASE maneuver is designed for maximizing the measurement capabilities of the instruments for atmospheric and ionosphere data profiles. The HOOK maneuver is aiming at improving the success in science payload recovery and in reducing the search and recovery efforts. To achieve these goals, a high performance and reliable flight control system is critical, that incorporates the throttling capability of the DVF hybrid rocket engine, which is one of the key development aspects of this study. For the numerical modeling of the internal ballistics of hybrid rocket combustion for flow analysis and design optimization, a multiphysics Navier-Stokes flow solver with finite-rate chemistry, real-fluid properties, turbulence model and radiative transfer model is employed for high resolution computations. This numerical model is also incorporated in analyzing the aerothermodynamics for high-speed ascend and reentry flights. A 6-DOF flight dynamics, navigation and control simulator is employed in assessing the overall performance of the vehicle based on the aerodynamics and propulsion data generated by the flow solver. Results of the numerical analyses are validated by measured data of ground and flight tests. KEYWORDS: Multifunction Sounding Rocket, DVF Hybrid Rocket Engine, Payload Recovery System, Trajectory and Flight Control, Aerothermodynamics. Copyright 2016 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

2 1. INTRODUCTION Sounding rocket system is a useful tool in many aspects of aerospace science and technology research and development since the early days of rocket experiments for space explorations. Mainly due to cost and system simplicity considerations, sounding rockets are traditionally spin stabilized without active controls such that the thrust profiles and rocket staging arrangements determine the nominally parabolic flight trajectories, which are subject to variations due to wind conditions and misalignments in the propulsion and aerodynamics subsystems. Flight guidance, navigation and control subsystems have been incorporated in modern sounding rocket developments to mediate the wind and misalignment effects during the early powered phase of the flight [1,2] that drastically reduces the dispersion in flight trajectories and improves the mission success. However, even with this method, payload recovery is not always guaranteed. To further enhance the benefit and outcome of the flight experiments using sounding rockets, fully active flight control subsystems can be employed such that the flight trajectories can be tailored for the purpose of achieving the main objectives of the experiments. Also, it is desirable that the payload can be flown back to an area near the launch site by the upper stage of the rocket after the experiment such that the payload recovery process can be much easily performed. To achieve all the aforementioned advanced capabilities, throttling authority of the propulsion system becomes critical in a fully active flight control system. Therefore, hybrid rocket propulsion is selected herein over its solid and liquid counterparts. This selection is based on technical capabilities and cost considerations. Hybrid rocket combustion research has received renewed attention in recent years due to its safety induced low development and operation cost and relatively benign environmental impact [3]. Although it is well know that the theoretical performance of hybrid propulsion systems can be higher than the conventional composite solid propellant systems, the existing hybrid systems fall short of their potential theoretical performance. This is mainly attributed to the diffusion flame combustion mechanism that reduces the overall combustion efficiency and affects the mixture ratio variations through the burn. Multiport designs, swirl injectors and liquefying paraffin grain are among the performance enhancing approaches for the hybrid rocket propulsion [3-7]. In the present line of research, innovative designs have been introduced in the past few years by employing mixing enhancement concepts with vortex generators and dual-vortical-flow (DVF) chambers [8,9]. In these studies, a high fidelity numerical model [10], with validations by hot-fire experiments, is employed in the study of the internal ballistics of hybrid combustion. These research efforts have resulted in a DVF hybrid rocket engine design that produces Isp of 292 sec using the N 2 O/HTPB system, which is close to 96% efficiency based on the theoretical thrust performance. Recent hot-fire tests of a subscale DVF hybrid engine with N 2 O/HDPE propellants has shown an averaged sea-level Isp of 280 sec. The results indicate that this model outperforms many other hybrid rocket designs and is getting closer to the performance of kerosene liquid engines. Based on these results, an advanced multifunction sounding rocket system is designed using a propulsion system based on the DVF hybrid rocket engine design and a fully active flight control system. Three basic flight trajectory designs are investigated in this study, namely the traditional standard parabolic trajectory, a TASE (Trajectory Augmented Science Experiments) maneuver and a HOOK (Homing Oriented Operation Kernel) maneuver. The TASE maneuver is designed for maximizing the measurement capabilities for the atmospheric and ionosphere data profiles. The

3 HOOK maneuver is aiming at greatly improving the success in science payload recovery by reducing risks in the search and recovery efforts. To achieve these goals, a high performance and reliable flight control system is also critical, that incorporates the throttling capability of the DVF hybrid rocket engine, which is one of the key development aspects of this study. 2. HYBRID ROCKET PROPULSION SYSTEM The multifunction sounding rocket system, AHR-I (Advanced Hybrid Rocket-1), proposed in this study consists of a two-stage hybrid rocket design with a fully active guidance, navigation and control system. The throttling capability of the hybrid rocket engine plus the attitude control ability provided by a thrust vectoring system enable AHR-I to perform trajectory following maneuvers tailored for the planned flight experiments. Table 1 shows some basic system parameters of a preliminary design. Table 1. AHR-I preliminary system parameters Launch Mass 1-Stage Total Impulse 2-Stage Total Impulse Rocket Size Construction 2-Stage Hybrid Sounding Rocket 1180 kg 162,309 kgf-sec 48,750 kgf-sec 9 m x 0.9 m Mainly Carbon Composite To deliver the required propulsion performance of the AHR-I system, thrust efficiency of traditional HTPB-based or HDPE-based hybrid rockets needs to be improved. For material availability and theoretical performance considerations, N 2 O/HTPB or N 2 O/HDPE is selected for the present development. This research effort was initiated by the authors in 2009, starting from a simple single port design with vacuum Isp = 200 sec. During the course of this research, two innovative designs were proposed, namely the Mixing-Enhancer (ME) design and the Dual-Vortical-Flow (DVF) design, to improve the vacuum Isp to 260 sec and 292 sec, respectively. Numerical modeling of the complex combustion physics and internal ballistics is accomplished by employing a high-fidelity multiphysics computational fluid dynamics (CFD) software, UNIC-UNS code [10], with parallel computing. Figures 1 and 2 show the computed flowfields for the ME and DVF hybrid rocket engines, respectively. The simulated thrust data of these engines were validated using measured data of hot-fire experiments [8,9]. Recent hot-fire tests of a 250 kgf DVF subscale N 2 O/HDPE engine have shown averaged sea-level Isp of 280 sec, which has provided a good technical guideline for scaling up designs to 3,500 kgf and 10,000 kgf thrust levels in the near future. A preliminary flight test of a 2-stage sounding rocket with 3,500 kgf and 1,000 kgf hybrid engines will be performed by the end of 2016 for technology demonstration. This flight test is also aiming at a launch altitude of over 100 km, which is an important step for testing the engine performance, stage-separation system function and the ignition of the second stage hybrid engine. Pyrotechnique and/or catalyst initiation ignition systems are considered for the ignition of the second stage. For demonstrating these technical features without system complications, the automated flight control system is deactivated for this flight. However, all attitude sensor data will be recorded and downloaded for the flight performance analysis purpose.

4 Fig. 1. Predicted internal flowfield of a hybrid rocket engine with mixing enhancers. Fig. 2. Predicted internal flowfield of a dual-vortical-flow chamber hybrid rocket engine. 3. FLIGHT GUIDANCE, NAVIGATION AND CONTROL SYSTEM To deliver all aforementioned performance and functions of the present multifunction hybrid sounding rocket system, the performance and reliability of the guidance, navigation and control (GNC) subsystem is very critical. Sensor accuracy, sampling rates, command delays, actuator delays and structural dynamics coupling are to be analyzed and well simulated in order to validate the proposed GNC system. The validation sequence includes pure software simulations, software-inthe-loop simulations and hardware-in-the-loop simulations. In addition, model-build-up flight test sequence is also required in the design process to validate and certify the GNC system step-by-step. This includes static tests, dynamic tests and various-scale flight tests. A single-stage flight test with flight control capability is also planned to be performed by the end of A thrust vector control (TVC) system is designed by taking advantage of the low slenderness ratio of the present DVF hybrid engine that is basically looked like a liquid rocket engine. All design practices for gimbaled liquid rocket engines can be adopted here. For simplicity, electric actuators are employed for the present TVC system. Figure 3 shows the sketch of a semi-distributed flight control avionics system designed for the present AHR-I sounding rocket system. This flight control system consists of a flight computer system with built-in redundancy, a data logger, IMU, accelerometers, GPS receiver, sensors (thermometers, barometer and strain gauges), thrust vectoring controller, telemetry and telecommunication (10W/434MHz and 7W/2.4GHz) and power system. System data rate is 100 Hz which applies to the flight control, sensors and telemetry data.

5 Fig. 3. Flight guidance, navigation and control system schematics. 4. REENTRY RECOVERY SYSTEM For the purpose of effective technology/instrument developments and cost considerations, payload recovery is in demand for flight experiments using sounding rockets. Traditionally, parachutes are used for landing the payload section on land or in the ocean after the experiment. Although the reentry speed of typical sounding rocket can be between 2 to 5 km/sec, which is not as high as that of orbital flights, the reentry heating the payload experiences can still damage its delicate instruments easily. A deployable membrane aeroshell reentry has been proposed and flight tested using a sounding rocket [11]. The results of this research have shown largely reduced reentry dynamic pressure and heating values. This concept is therefore adopted in the development of the present sounding rocket payload recovery system. The aeroshell made of heat-resistant fabric is tailored and folded in a space around the lower part of the upper stage under the payload section. A small high-pressure nitrogen gas canister is installed and commanded to inflate the aeroshell in the descending phase of the flight above 100 km altitude where air resistance is still small for safe deployment of the aeroshell. The pressure in the aeroshell is regulated such that appropriate structural strength of the aeroshell can be maintained during the reentry flight. A flight test for the reentry of a model aeroshell will be performed in early 2017 to take some heating the pressure data on the forward-facing surface of the aeroshell. The measured data are very useful for validating the numerical simulation results and for guiding the design of the final model of the reentry system. The present CFD method is applied in the design and simulations of an aeroshell configuration with 60-degree half angle. The computed aerodynamics data are then used in a 6-DOF trajectory simulation software for assessing the overall performance of the present aeroshell design. Figure 4 shows the calculated flow fields for Mach 2, Mach 4 and Mach 8 flight speeds. Figure 5 shows the comparisons of flight performance between traditional reentry capsule and the present aeroshell reentry system. It can be clearly seen the effects of aero-breaking the dynamic pressure reduction of the aeroshell, which indicates a great reduction in the maximum heat load in the reentry flight. Figure 6 illustrates the launch operation sequence and the launch performance of the preliminary design, with payload recovery capability. The proposed approach of flying the upper stage with payload back to the vicinity of the launch site, which is designed to be around 5 km (within visible

6 range) off the seashore of the south east part of Taiwan, the chance for success of payload recovery can therefore be largely increased. The aeroshell reentry system technology developed in the present study can be applied to orbital flight conditions in the next phase. Much higher head load can be expected for orbital reentry. Thus, material selection and aeroshell configuration optimization are critical areas for the design of such system. For rarefied flow regimes, direct simulation Monte-Carlo (DSMC) method can be employed for better flow solutions with more accurate shock resolutions and surface heat flux predictions. (a) Mach 2 (b) Mach 4 (c) Mach 8 Fig. 4. Aeroshell reentry aerodynamics simulations (Mach number contours). Fig. 5. Flight trajectory and dynamic pressure comparisions.

7 Fig. 6. Flight operation system and typical flight sequence. 5. TAILORED FLIGHT TRAJECTORIES For science and technology flight experiments, it is highly desirable that the flight trajectories can be well predicted in order to maximize the outcome of the experiment. In the case of ionosphere science experiments, for example, a lot more scientific information may be revealed if horizontal profiles can be measured in one flight. In other examples, flight path designs can be very important for the missions such as reentry and/or hypersonic flight experiments. To fulfill the requirements of this type of experiments, a TASE (Trajectory Augmented Science Experiments) maneuver function will be implemented in the proposed multifunction sounding rocket flight control system.

8 It is particularly true for seashore launches of sounding rockets (typical in Taiwan) that payload recovery effort becomes very challenging since the landing location is far away and the sea and weather conditions can often be very unforgiving. These payload recovery difficulties can be greatly reduced if the upper stage can be designed to carry the payload through a circular flight path after the experiment and fly back to the vicinity of the launch site then perform the reentry and landing in the sea area close to the shore, e.g. 5 km from the coast that is within visible range. To do this, a HOOK (Homing Oriented Operation Kernel) maneuver function will also be implemented in the present sounding rocket system. Figure 7 illustrates 3 tailored flight trajectories that include the traditional parabolic, TASE and HOOK flight paths. For the HOOK maneuver, proper amount of propellants is reserved after the mission to perform the turning around of the rocket, to fire the hybrid engine to establish the return flight trajectory, and to adjust the flight path aiming at the landing zone before the deployment of the aeroshell for reentry. Propellant management for performing this maneuver is a critical issue in the design process. Some flight tests are required to fine tune these control parameters. Fig. 7. Comparisons of different flight trajectories. 6. CONCLUSIONS In this study, we have described the outline of a development plan of an advanced multifunction sounding rocket system using high-performance dual-vortical-flow hybrid rocket propulsion with a full active flight guidance, navigation and control system. The flight control system features dual flight computers for redundancy implementation for improved reliability. The dual-vortical-flow hybrid rocket engines incorporated in the present study delivers high thrust performance (vacuum Isp of 292 sec) with high combustion efficiency due to its improved internal ballistics design that enhances the mixing and combustion processes in the combustion chamber. This hybrid propulsion system is designed for N 2 O/HTPB or N 2 O/PE propellant combination. Three basic flight trajectory designs are proposed in this study, namely traditional standard parabolic trajectory, TASE (Trajectory Augmented Science Experiments) maneuver and HOOK (Homing Oriented Operation Kernel) maneuver. The TASE maneuver is designed for maximizing the measurement capabilities of the instruments for atmospheric and ionosphere data profiles. The

9 HOOK maneuver is aiming at improving the success in science payload recovery and in reducing the search and recovery efforts. ACKNOWLEDGEMENTS This work is supported by the research grant of the Ministry of Science and Technology and the National Space Organization of the National Applied Research Laboratory, Taiwan. The computational resources are provided by the National Center for High-performance Computing of the National Applied Research Laboratory. REFERENCES [1] Ljunge, L.F., Sounding Rocket Guidance, Navigation and Control and Anticipated Next Generation Developments, 29 th International Symposium on Space Technology and Science, 2013-m-08, June 2-9, 2013, Nagoya-Aichi, Japan. [2] Ljunge, L.F. and HALL, L., S19 guidance of the Black Brant X sounding rocket, Journal of Guidance, Control, and Dynamics, Vol. 7, No. 2 (1984), pp [3] Chiaverini, M.I. and Kuo, K.K. (Editors), Fundamentals of Hybrid Rocket Combustion and Propulsion, Vol. 218, Progress in Astronautics and Aeronautics, [4] Story, G., Zoladz, T., Arves, J., Kearney, D., Hybrid Propulsion Demonstration Program 250 K Hybrid Motor, AIAA , 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Huntsville, AL, [5] Karabeyoglu, A., Zilliac, G., Cantwell, B.J., DeZilwa, S., and Castellucci, P., Scale-up Tests of High Regression Rate Paraffin-Based Hybrid Rocket Fuels, J. of Propulsion and Power, Voll 20, No. 6, p , [6] Karabeyoglu, A., Stevens, J., Geyzel, D., Cantwell, B., and Micheletti, D., High Performance Hybrid Upper Stage Motor, 47 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, July 14-17, 2011, San Jose, CA. [7] Chandler, A., Jens, E., Cantwell, B., Hubbard, G.S., Visualization of the Liquid Layer Combustion of Paraffin Fuel for Hybrid Rocket Applications, AIAA , 48 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, July 30-Aug. 1, 2012, Atlanta, GA. [8] Chen, Y.S., Chou, T.H., and Wu, J.S., N2O-HTPB Hybrid Rocket Combustion Modeling with Mixing Enhancement Designs, AIAA , 49 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, July 14-17, 2013, San Jose, CA. [9] Chen, Y.S., Lin, J.W., Chou, T.H., and Wu, J.S., HTPB Hybrid Propulsion with Multiple Vortical-Flow Chamber Designs, AIAA , 50 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, July 28-30, 2014, Cleveland, OH. [10] Chen, Y.S., Chou, T.H., Gu, B.R., Wu, J.S., Wu, B., Lian, Y.Y., and Yang, L., Multiphysics Simulations of Hybrid Rocket Combustion, Computers & Fluids, Vol. 45, p. 29, [11] Yamada, K., Suzuki, K., Abe, T., Nagata, Y., Imamura, O., Akita, D., Takahashi, Y., Honna, N., Watanabe, Y., Iino, T., Sasaki, K., Atmospheric-entry Flight Test of Deployable Membrane Aeroshell Using S-310 Sounding Rocket, 29 th International Symposium on Space Technology and Science, 2013-m-10, June 2-9, 2013, Nagoya-Aichi, Japan.