Space Future - Flight Demonstration and a Concept for Readiness of Fully Reusable Rocket Vehicles

Abstract

A fully reusable rocket vehicle is proposed to demonstrate good operability characteristics both on the ground and in flight. For achieving technical readiness for future space transportation systems, design considerations not only for higher performance-related issues but also those for good operability are needed. The proposed vehicle is to be used as a sounding rocket and has the capabilities of ballistic flight, returning to the launch site, and landing vertically making use of clustered liquid hydrogen rocket engines. Before initiating the development of this type of reusable rocket, a small test vehicle with a liquid hydrogen rocket engine was built and flight tested. A demonstration of vertical landing and exercise of turnaround operation for repeated flights are the major objectives of the test vehicle. Two series of flight tests were performed in 1999 and 2001 and the flight test operation provided repeated flight environment and many lessons valuable for designing the fully reusable rocket vehicle.

1. Introduction

A number of performance-related innovations are needed for the future space transportation systems such as fully reusable launch vehicles. The high performance engines and nozzles, super-light weight materials and structures are the necessary technical items to be ready to build these future vehicles. In addition to these technical challenges, it is required that the flight and ground operations of these advanced vehicles should be closer to those of aircraft. It is believed to be a key to an order-of-magnitude cost reduction of the transportation between ground and low earth orbit, which is a major element of the whole space infrastructure in the future. Current expendable launch systems prevent it from achieving the easy access to space because of its ammunition-type of operation. The discussion here concerns the space transportation architecture that would be achieved by the existing technology basis and for the time period of coming twenty to thirty years.

Current space transportation demands and markets such as broadcasting, communication, navigation, space station's logistics, and so on, will be covered by existing launch vehicles. Then what should be our next goal? For the future vehicles by which a true reusability is beneficial, a tremendous cost reduction is needed and expected. To what extent reduction we need, then? Although many ideas for these future demands have been figured out and proposed, some of the synthesis of realistic and economically viable transportation demands tell that tourism by general public and a construction of huge space structure such as solar power satellite ( SPS) have been studied extensively and quantitatively1. A good "reusability" of the launch vehicle is absolutely necessary for such a goal. Studies for transportation systems must be made not only from view point of technical aspects but also from view points of economical and demand-oriented aspects, and acceptance by society. In such future demands, it is concluded that two-order-of-magnitude cost reduction is necessary, and that one-order-of-magnitude reduction and less is not enough. A mass transportation and tremendously frequent flight by many vehicles' fleet must be established as compared with the existing launch systems. In such a way, benefit of the true reusability or aircraft-type operation of the vehicle is given for such a future demands. In other words, the demands of today for coming ten to fifteen years does not need such a new characteristics of the true reusability as presented in Fig.1

To "change launch culture" should be a key to the future system of space launch in such a totally different world. It means that the current launch culture by expendable launch system is not that for "transportation" but an extension of the warhead launch or ammunition2. These characteristics result in a few launch per year, thousands of standing army for each launch preparation, long waiting time and delay, each of which automatically costs a lot. Orders-of-magnitude cost reduction is possible only by changing the culture. A system like an aircraft only will achieve the goal of reduction presented above by its simplified operation, flight on demand, quick turnaround and continuous intact abort capabilities. These features are so called aircraft-type operation in comparison with the current existing launch systems. Not only technical studies to be ready for the advanced vehicles, but also identifying the many necessary conditions will be quite meaningful and must be essential to figure out what our goal is, and how we are ready to achieve this goal.

Taking these issues into consideration, a conceptual vehicle system of the reusable sounding rocket is proposed, and this preliminary design consideration will point out that these necessary and essential aspects would be common among these types of fully reusable space transportation systems to be built. In addition to these design considerations, a small test vehicle was built and flight-tested in order to experience how the reusable vehicles should be designed and operated. Since the capability of the vehicle is limited, it focuses on the landing flight characteristics and turnaround operation of the vehicle of this kind. An importance of these considerations is to be understood by demonstrating the benefit of the repeated flight environment for the new technology research and it is one of the keys to accelerate further studies.

2. Reusable Sounding Rocket

Referring to the discussion made previously, a rocket vehicle is preliminarily designed3,4. A goal of the vehicle is first to achieve the fully reusable vehicle with enhanced operability, which will demonstrate the benefit of reusability. At the same time, the rocket vehicle is used as a sounding rocket. An easy access to the flight opportunity is quite important for those who wish to use it such as the astrophysicist and researchers of atmosphere. The micro-gravity community is also a potential user of the vehicle. By enhancing the flight operability, low cost operation of the vehicle will give a good opportunity for these researchers or users, which means the frequent use of the vehicle is expected. Since the development of the orbital vehicle is a huge business, starting from a small but good reusability vehicle would be one of the effective ways to the goal, and the flight opportunity by the vehicle would potentially be beneficial to users. These are the background idea for the present study. In addition, new technologies necessary for the future vehicle like SSTO, such as an altitude compensation nozzle and new materials, will require an in-flight demonstration, because it is very difficult to certify these new capabilities by the ground based facilities. Thus the test vehicle will be also beneficial to technical studies.

As for the rocket performance, the proposed vehicle has a ballistic flight capability to the altitude up to 300km, and safely returns to the launch site. Finally it lands vertically and provides the safe return of the payload without any difficulty in its recovery like conventional sounding rockets. As it is presented in the previous section, a quick turnaround capability will give a repeated use by the user or several times of the flight in one launch campaign would be made possible. A design goal of the turnaround interval should be a "daily flight". For a higher lift-to-drag ratio vehicle such as winged vehicles which take relatively shallow entry into atmosphere, it is difficult to come back to the launch site after the ballistic flight without propulsion system for the returning cruise flight. Preparing both launch and landing site in many hundreds or thousand-kilometer distance apart will not be beneficial for the effective reusability lessons. Since one of the objectives of the present system is to show the good reusability, a flight operation by a single ground facility is to be a major necessary condition. Thus a vertical-lander-type system is chosen in the present study. A flight nearly vertical will make it possible, and the vehicle with relatively low lift-to-drag ratio will be capable of this type of flight. A preliminary flight analysis shows that a base-entry vehicle with 0.3 of lift-to-drag ratio performs aerodynamic returning flight to the launch site. Then it restarts the engines for final approach and landing flight without an attitude turnover.

Fig.1 Proposed Reusable Sounding Rocket


VEHICLE

LENGTH (m)	4.4
BODY DIA. (m)	2.2
LIFT-OFF WT. (kg)	3800
LANDING WT. (kg)	1460
PROPELLANT WT. (kg)	2400
PAYLOAD WT. (kg)	100
MAX ALT. (km)	300

ENGINE

PROPELLANT	LOX/ LH2
Isp	350/400
THRUST (ton)	1.7 (S.L.)
DRAY WT. (kg)	70
THROTTLING (%)	100-30

Table 1 Dimensions & Engine Specifications of Reusable Sounding Rocket

Figure 1 presents the resulted vehicle. Its propulsion system is composed of four liquid hydrogen / liquid oxygen engines, and it is lifted-off vertically as a conventional rocket, then the vehicle reaches to an altitude of 300 km. After falling down to the atmosphere, it is aerodynamically decelerated and lands vertically at the same place it is launched. The payload carried to the altitude is 100 kg which is equivalent to the current sounding rocket of ISAS (Institute of Space and Astronautical Science ). Table 1 summarizes its dimensions and engine specifications. Considerations for the design of continuous safe flight abort and effective ground operations were made extensively. Study results of operational and safety aspect of hydrogen aircrat5, Shuttle flight abort6, and lessons learned by DC-X and XA7, were taken into the present preliminary design considerations.

3. Exercises by Flight Test Vehicle and Lessons Learned

3.1 Test Vehicle Overview

A small test vehicle was built and flight-tested in order to be partly ready for the design of the vehicle proposed above. This is Reusable Vehicle Testing (RVT) lessons campaign, which will give us an opportunity to exercise new design and operational aspect peculiar for the reusable rocket vehicle. Following two primary objectives are addressed for this RVT exercise out of many important design issues presented in the previous sections; 1) Design and operational lessons in repeated flight of rocket and its turnaround, and 2) Vertical landing of rocket propelled vehicle. For designing the fully reusable rocket vehicle, these two topics are ones of the essential characteristics. It is necessary to experience these in-flight and on-the-ground operations for the design and construction of the reusable vehicle. In addition, making maximum use of the repeated flight environment, the vehicle develops as long as the flight test going. New flight test articles are to be added for the succeeding in-flight demonstration. This is one of the important characteristics of "reusable" vehicle in contrast with the expendable vehicle development.

Fig.2 RVT Test Vehicle


VEHICLE DRY WEIGHT	320 kg
HEIGHT	3.0 m
ENGINE	LOX/ LH2 (PRESSURE FED)
THRUST LEVEL	280 - 400 kg (THROTTLING BY LOX FEED)
PROPELLANT	LH2: 100 l, LOX: 42 l (MAXIMUM)
FLIGHT TIME	20 sec. (MAXIMUM)
ATTITUDE CONTROL	GN2 RCS
NAV. SOURCE	IMU + LASER ALTIMETER
LANDING GEARS	FIXED + DAMPER
TURNAROUND	2 FLIGHTS/DAY (TARGET)

Table 2 Specification of RVT Test Vehicle

Referring to the performance oriented requirements of the future vehicle, liquid hydrogen is an only fuel that gives promise to the higher performance rocket such as SSTO. Regarding with the ground operation of the vehicle, a cryogenic fuel will impose an essential constraint for its operational and safety aspects. Therefore a slightly modified engine based on the ISAS 's existing LOX/ LH2 small rocket engine is used for the vehicle. It is a regeneratively cooled engine with a continuous throttling capability for making vertical landing possible. Although this engine is not originally designed for its reusability, it was estimated that at least 30 times of firing was possible from view point of stress level of its combustion chamber material. The vehicle's fueling subsystems are designed so as to achieve the repeated flight easily and quick turnaround. The vehicle's dimensions and specifications are presented in Table 2 and outlook of the vehicle is presented in Fig. 2.

The vehicle's flight navigation, guidance and control subsystems are built based on the conventional Inertial Measurement Unit (IMU) and a laser altimeter. The attitude of the vehicle is controlled by three-axis RCS by cold nitrogen gas jet system placed on the top of the vehicle. The test vehicle has four-leg landing gear which make it possible to eliminate any launch support nor launch tower. The design landing speed is less than 1 m/sec. and the maximum allowable is 3 m/sec. from structural and damper stroke limitations. The final landing guidance concept is to let the descending speed follow the predetermined line with respect to the altitude as described later in the flight result. At the instance of landing, the engine is to be cut-off automatically with respect to the altitude and altimeter-derived sink rate measurement, which means no landing sensors nor touch sensor to the ground is equipped with.

The engine is pressure-fed system for the sake of simplicity. The fuel and oxidizer tanks are pressurized by helium bottles on both sides of the vehicle, as shown in Fig. 2. For the first flight vehicle, engine throttling is made by controlling the flow rate of the liquid oxygen alone for simplicity, too. Therefore the oxidizer / fuel mixture ratio is varied between 3 to 7 with respect to the level of the thrust. To chill down the engine before ignition, the venting of the fuel is done and freely disposed into the air horizontally from the top of the vehicle. It also helps eliminate the necessary ground venting lines, quick disconnect couplings and associated equipment. Liquid oxygen for chilling down is vented downward from the vehicle and it is dumped out about five meters apart from the vehicle through a pipe on the ground in order to separate the freely dumped fuel and oxidizer. The telemetry subsystem sends the measured status of various on-board subsystems, and it is monitored by ground operators. Emergency commands are to be sent when the anomalous status was found. Then the vehicle's flight mode is to be transferred to the emergency landing or to stop the engine when the range safety issues would occur.

Prior to the flight testing, two series of engine firing test were carried out. The engine's ignition and cut-off transient, static characteristics and dynamic throttling responses were characterized and identified. In order to achieve the safe and soft landing, the throttling characteristics of the engine is required in both static and dynamic manner. The engine is a part of the guidance / control plant of this flight test system. Referring to the landing flight analysis, 1Hz and higher response of the thrust control requirement was imposed to the dynamic throttling characteristics of the engine. Due to the simplicity of the pressure-fed system, it was not very difficult to satisfy this request by choosing the appropriate proportional valve with electrically driven actuator. The resulted frequency response of the engine thrust was characterized as presented in Fig. 3 which is sufficient to conduct the flight, and these characteristics were used for the guidance and control analyses. The estimated dynamic responses using simulation model was constructed by modifying a numerical simulation tool developed for the pump-fed system8. During these firing tests, engine-firing-oriented mechanical environment to the vehicle structure and on-board instruments are also identified. At the last stage of these firing tests, the vehicle stand-alone firing was made. By the test, the overall functions of the on-board subsystems were qualified, and the mechanical and thermal environments were characterized particularly from view point of the ground effect at its vertical take-off and landing.

A planned flight sequence is presented in Fig.4. On each mode of the flight and on the ground, an on-board status monitoring system is activated. Unlike other conventional expendable vehicles, the test vehicle has an intact abort capability which means that the abort at any time of the flight, and transition to the safe landing are executable. This abort capability is also one of the new characters of the reusable vehicle. The on-board monitoring system checks the status of the engine, pressurized tanks, RCS and on-board computer itself. The system automatically estimates the predicted states and allowable band of several important state variables, such as engine combustion pressure and temperatures of critical parts of the engine. These in-flight estimates were done with respect to the level of the engine throttling in which the transient responses are also taken into account, because its dynamic throttling range is wide. Once one of these monitored variables would travel out of the allowable tolerance, it automatically transitions to the safe landing mode.

Fig.3 Dynamic Throttling Characteristics

Fig.4 Flight Sequence of the RVT#1 Vertical take-off & Landing Flight

3.2 First Flight Test (RVT#1) Results

After completion of the preflight tests and synthesis, the flight test was conducted in March 1999 at ISAS 's Noshiro Testing Center (NTC). Following the functional tests and engine's static firing tests in tied-up configuration on the ground, two flights were carried out. In the second flight as shown in Fig.5, a full vehicle motion as shown in Fig.5 was performed. The flight time was 11.5 seconds and the maximum altitude was 4 meters. These two flights were conducted within two days. The vehicle performed the flights as planned, and the vehicle's propulsion and navigation/guidance/control systems worked as expected. The horizontal conversion of the vehicle's position in the flight was 4 m. For the landing, The vehicle kept almost constant sinking rate and engine throttling was properly controlled with respect to the guidance command. Figure 6 shows the history of the resulted landing speed with respect to the altitude. It shows that the landing guidance made the vehicle follow the predetermined lines as described earlier. In the present flight, the landing guidance target was set to be almost constant landing speed as shown in the figure. The landing impact was smaller than anticipated, and the dampers of four legs almost did not stroke. The engine was cut off almost at the instance of the ground impact as planned. The free venting of the fuel and oxidizer before the flight for chilling down the engine for ignition showed in good shape. The final automatic chill down sequence lasted 12.5 sec for liquid oxygen and 5.5 sec for liquid hydrogen, respectively, they were cut off and 0.5 sec after the engine's ignition. The flow rate for chill-down is 1.4 liter/sec for liquid hydrogen and 0.22 liter/sec for liquid oxygen, respectively. According to the ground wind direction and speed, the vented hydrogen gas stays around or flows out of the vehicle. In some cases it slowly burned up by the engine's exhaust at its ignition, but in other cases not. In every engine ignition and firing during the test including the ground firing test throughout the campaign, there was no impact to the vehicle and flight at all.

Fig.5 RVT Test Vehicle in Flight

Fig.6 Landing History in the Second Flight

Two flight operations and the ground firing tests in the campaign were made in a daily basis as presented earlier. The turnaround of the vehicle was conducted as shown in Fig.7. The flight operation is initiated by positioning the vehicle to the lift-off point from the hanger. After completion of the electrical functional tests, the pressurization of the RCS and accumulators of propulsion subsystem are made. Following the liquid oxygen, liquid hydrogen is fueled. These preparation works took two hours from the beginning by three crews worked neighboring to the vehicle for making propulsion subsystems ready. After the evacuation of the engine crews to the safety distance, the fuel and oxidizer tanks were pressurized remotely. Initializing the navigation system, countdown operation starts the engine to chill down. The propulsion subsystem's valve control authority is handed over to the vehicle's onboard computer at -30 seconds and automatic ignition sequence starts.. Three seconds before the take-off, the engine is started at the minimum of its thrust level. Then the vehicle performs take-off and landing flight. Immediately after engine's cut-off at landing, purging to prevent the engine from freezing is started automatically from the on-board helium accumulators. After the flight, the status of all the subsystem is checked remotely via telemetry and visual monitors, and the fuel and oxidizer tanks are depressurized to the safety levels by sending the commands. The engine crew comes access to the vehicle and reconnect the ground support lines for control signal and power supply. Then the valve control authority returns to the ground. It takes about 5 minutes for the post-landing activities to make the vehicle safe. The residual fuel dumping is done by connecting the vent lines and stuck apart from the vehicle. These works are also done by three members of engine crew. After 30 minutes of fuel dumping is completed, the vehicle becomes totally safe, and the deactivation of the vehicle and a series of post flight inspection are started. It takes within 4 hours to conduct all the activity from the beginning. These operations were all done in safe condition. As for the turnaround and readiness to the next flight, a visual inspection of the engine and a leak detection test of the propulsion subsystem were done by sealing the nozzle exit after each flight. As a result, the test system including the vehicle itself and the ground support functioned as planned and expected, and the daily turnaround was achieved.

Fig.7 Flight Turnaround Operations

Fig.8 RVT#2 Primary Propulsion Subsystem

3.3 Second Flight Test (RVT#2)

After the success of the first flight testing, primarily aiming at an extension of its flight envelope, many parts of the vehicle was replaced and modified. The vehicle's core structure and landing gears were unchanged, however the engine's combustion chamber, and fuel and oxidizer tanks were replaced for higher level of thrust. The new combustion chamber was manufactured based on the life-controlled and stress-relaxation design particularly to the regenerative cooling passage. The RTK (Real-Time-Kinematics) GPS (Global-Positioning-System) was employed for augmenting navigation capability in addition to the previously employed inertial navigation and laser altimeter subsystems. Primary reason for the augmentation is due to higher altitude flight, however another reason is to avoid the shielding effect of altimeter's laser beam by freely vented Hydrogen and Oxygen cloud shortly after the lift off as presented before. To be ready for the further extension of the flight envelope, airframe was installed for managing the aerodynamic forces and moments. Even for a low subsonic flight speed of the vehicle, many aerodynamic effects such as a side wind effect, an engine exhaust interaction with the ground and free stream and so on were carefully examined for the coming extension of the flight regime9. The airframe was made up with shells of glass fiber reinforced plastic. The base surface of the airframe was covered by flexible thermal insulation to prevent the airframe from heating by ground radiation made by exhaust plume. Covering the vehicle's internal structure and propulsion subsystem by airframe will impose a serious concern about the leakage of the hydrogen, which may result in a fatal consequence. Therefore a gaseous hydrogen leakage detector was installed inside the vehicle and the leakage signal is to be monitored throughout the test operation including the in-flight detection. This leakage sensor signal is added to the in-flight monitoring subsystem presented above, and once the leakage was detected the vehicle was to be transitioned to the emergency landing mode. For the propulsion subsystem, a new proportional valve was added to the fuel supply line as well as to the oxygen line, in order to keep a constant fuel / oxidizer mixture ratio. Two Helium bottles were added for pressurization of the propulsion subsystem. Figure 8 shows the schematic diagram of the propulsion subsystem, where subsystem's components and external interfaces are presented. Prior to the flight, two series of ground firing tests were performed and both engine's static and dynamic thrust control performance were characterized as it was done before the first flight test. At the firing tests, most of the vehicle's subsystems were put onboard and the engine derived mechanical and thermal environments were also characterized. These qualification processes are the same as was done in the first flight test vehicle. As a result the vehicle weight became almost double the first flight test configuration, however owing to the engine thrust augmentation, the flight acceleration of the vehicle is about 1.3G compared to 1.15G of the first flight test vehicle. For the better attitude control characteristics, the control moment by the reaction control subsystem was also augmented accordingly. Onboard vehicle management and health monitoring software was similar to those of the first flight test, however many parameters were tuned and the structure of the software refined based on the lessons learned in the first flight test and further studies.

In June 2001, the second flight test were performed at NTC of ISAS as well as the previous flight test (Fig.9). Three times of flight within three and half days were made after a ground-firing test for the final qualification of the whole test system and operational readiness. The ground support subsystems were totally the same as those of the first flight test campaign except the ground station of RTK-GPS and its monitoring system. Each lift-off and landing flight exhibited fairly good results as expected. In the last flight the vehicle reached to the altitude of 22m and flight time was about 12 seconds. The guidance laws and engine-cut-off procedures at its landing were in the same way as well as those of the first test. These three flight trajectories are presented in Fig.10. Both IMU-derived, laser-altimeter-derived, and RTK-GPS-derived altitude histories are presented in each trajectory result. The onborad navigation signal in the guidance loop in the first flight was made by only IMU and laser altimeter. The GPS navigation source was out of the loop in order to check the quality of the signal. The second and third flight employed the GPS signal instead of laser altimeter, and exhibited satisfactorily good navigation and guidance performance. In each flight, the landing speed was kept constant by almost 0.5 m/sec, and all landing behavior showed smooth motion, which resulted almost no stroke of the landing damper. Rest of the vehicle's subsystems worked out totally as expected including hydrogen leakage detection. As presented previously the vehicle employed the free venting for chill down operation before ignition. A considerable amount of gaseous hydrogen is exhausted out of the vehicle, and the flights were made under slight and almost no ground wind conditions, however there were no hazardous problem occurred by the airframe installation. As a result, the test campaign was made in the way of "flight envelope extension" by adding new test articles. Turn around operation was also made smooth, which resulted in daily flight operation again without any serious issues.

Fig.9 Second Flight Test (RVT#2)