Preliminary Design of the Hybrid Air-launching Rocket for Nanosat

Jae-Woo Lee, Kyung-Ho Noh and Yung-Hwan Byun Department Aerospace Information Engineering, Konkuk University, Korea

Bong-Kyo Park Hanwha Corporation jwlee@konkuk.ac.kr

Abstract

Air-Launching is an effective method that can launch the 'Nanosat' for low launching cost. In this study, the preliminary design of the air-launching micro rocket was performed. Mission and trajectory optimization was performed and the propulsion system was designed based on the mission design. A supersonic air launching rocket with total weight of 830.95 kg, the payload weight of 3.56 kg and the propulsion system length of 5.89 m has been designed. 1. Introduction

Recently, the demand for the small satellites is growing and many countries which hold advanced technology in the space launcher development, concentrate on the development of the rockets for the small satellites. For the development of very small satellites, less than 10 kg of weight, the space technology which applies the MEMS(Micro-Electro Mechanical Systems) is necessary[1, 2].

Very small satellites can be classified into two categories; Nano-satellites, which weigh around 1 - 10kg, and pico-satellites, which weigh around 0.1 - 1kg[3]. Several nanosats have been developed till now: the QQW1 of 4 kg weight, the ODERACS series satellites of 5 kg weight, and 3 kg satellites like the SPUTNIK-40 and the TUBSAT-N[1]. Nanosat is regarded as the future-oriented new technology which can overcome the limits of the small and micro satellites. Key factors when designing or selecting a launch vehicle are the launch cost and the launch capacity, i.e. the weight of the satellite to be carried. Because the launch cost per unit mass will grow as the weight of the satellite become smaller, either several nanosats must be launched together or, the nanosat launched with the large satellite. Hence the launching time and the operation of the satellite are limited. Therefore, new launching method can launch the

nanosat individually with low launching cost. 'air-launching' can be a solution.

By implementing air-launching, there would be no restrictions on the launch sites, the launch angle and the launch direction. This can be a very strong point especially to the countries where the satellite launching is very difficult owing to the geographical reasons, like Korea. Moreover, "air-launch" is a very economical way of launching satellites compared with the ground-launch, because it can utilize the high initial launching speed from the mother plane, and the improved thrust efficiency resulted from low dynamic pressure and big nozzle expansion ratio at high altitude[4, 5].

Among the air-launching rocket system developed until now, there are NOTS and Cabel. NOTS, the first air-launching rocket, was launched by a Douglas F-4D Skyray airplane. As the vehicle was designed for maximum simplicity, it featured no moving parts. All launches apparently failed, most due to structural failures. The first and third orbital attempts could have possibly orbited their payloads, as tracking stations picked up signals, which could have originated from the payload. After 10 launches the program was discontinued and a improved vehicle, called Caleb, was developed[6]. The Pegasus and the Burlak are more recent air-launchers. In the M-3S II User Guide, ISAS of Japan also announced its intention to develop air-launching rockets for the small satellite using M-V rocket[7].

Pegasus is a small commercial launch vehicle developed by Orbital Science Corporation. It is a three-stage, solid-propellant, inertially guided, all-composite winged space booster. The Pegasus has been carried aloft by a specially modified Lockheed L-1011 carrier aircraft to level flight launch conditions of approximately 11.9 km altitude and Mach number of 0.8. Pegasus follows a nearly vacuum optimized lifting ascent trajectory to orbit, carrying 275 kg payloads to 480 km polar orbits as well as proportional payloads to

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other altitudes and inclinations or suborbital trajectories[8].

Burlak, a Russian Pegasus, uses a liquid propellant and is launched at 13.5km with a supersonic speed, M=1.7 from the Tu-160 bomber[5]. As the launching speed from the mother plane increases, the weight and launch cost can be reduced, therefore the supersonically launching Burlak is more advantageous than the Pegasus, launched at subsonic speed. Various air-launchers are shown at Fig. 1.

Figure 1. Air-launching rockets In this study, the design process of the three-stage

air-launching rocket system shall be defined first, and the mission design, the conceptual and preliminary design and system optimization will be performed. 2. Air-Launching Rocket Design Process

For the air-launching rocket design, like the aircraft design, many technical areas like aerodynamics, propulsion analysis are involved from the concept exploration stage, hence an integrated design consideration among all the related engineering fields is crucial to acquire the best design solution with given constraints in order to reduce the design changes or resulting development cost.

When developing a system, with which many engineering disciplines are highly coupled like rocket system design, the classical separate single discipline design approach is difficult to apply, because most design variables are coupled with many engineering phenomena. To resolve the problems which occur during the design optimization procedure, Multidisciplinary Design Optimization (MDO) techniques, which can handle complex engineering problems with many design variables and constraints from various technical disciplines, are required[9].

Figure 2. Air-launching rocket design process

With given design requirements, rocket mission is defined first. Then the propulsion system including the rocket motor, the oxidizer feeding system, and the fuel/oxidizer control system is designed. Through stability and control analysis and aerodynamic analysis, the wing and tails are designed. Subsystem are defined and designed. The material properties of the major structural components and subsystems are defined hence, approximate weight of the rocket system is estimated. The optimized design can not be obtained at one design cycle, several design iterations are required. The rocket design process is shown at Fig. 2.

In this study, the mission design, the trajectory design and optimization, propulsion system design, wing design, and the subsystem design will be performed.

3. Design of the Hybrid Air-launching Rocket for nanosat 3.1. Mission Design and Optimization

The mission design process of the air-launching rocket system can be divided into several parts. Based on the given design conditions and mission

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requirements, the characteristics of the propulsion and basic aerodynamic parameters are determined. The required velocity to launch the payload to the pre-specified orbit, the mass ratio of each stage, and the structural coefficient are estimated. From the simple empirical equations, the total required velocity of the launcher is calculated by considering the gravity, the drag, and the thrust loss. The required total velocity is distributed to each stage in order to have the minimum total weight. From the trajectory analysis and optimization, a mission profile which can minimize the total rocket weight, is derived. This process is repeated until the design constraints are all satisfied.

For the air-launching rocket design study, F-4E, which has the best weapon load capacity in ROK Air Force inventory, is selected as the mother plane. The lowest possible latitude location while not going through the main land of Japan, must be selected as the air-launching site, in order to maximize the earth rotation effect. The mother plane will take off at the Jeju airport (the longitude 126°29′ 42″ , the latitude of 33°30′ 29″ ), and the launching location is the longitude 125°, the latitude of 30°(Refer Fig. 3). The total mission radius is approximately 682 km.

For the mission analysis of the mother plane, the mission profile is divided into 16 mission segments and the "MISS V 2.01"[10], a general aircraft mission analysis program, is utilized. Total required fuel weight is 12000 lb, 4670 lb less than the maximum fuel weight of the F-4E. Therefore, it is demonstrated that the flight mission of the mother plane to the launch location is possible.

Launch location

Figure 3. Air-launching location

From the variables determined so far, the total required velocity to launch the rocket to the specified orbit is calculated using the empirical equations. Drag

is influenced only on the first stage, and the velocity loss due to the atmospheric pressure is neglected, considering the low atmospheric pressure and high thrust-to-weight ratio during the second and third stages. In addition, velocity gain from the rotation of the earth, and the air-launching initial velocity gain are included in the determination of the total required velocity of the launcher. The obtained required velocity is 8966 m/s. The total velocity is distributed to each stage of the rocket in order to minimize the rocket weight. Table 1 shows the result.

Table 1. Velocity distribution

Stage Structure(kg) Propellant(kg) Payload(kg) Total(kg)

3rd 2.95 10.45 3.5 16.90

2nd 16.69 62.77 17.43 96.36

Fairing 10.0 • • •

1st 152.15 572.36 106.36 830.87

3.2. Trajectory Optimization

The trajectory analysis and optimization code is developed using the simple equations of motion, which were derived by assuming the rocket as a point mass and the earth as a perfect sphere[11]. To simulated the lift produced from the wing at first stage of the rocket the aerodynamic model[12] of Pegasus launcher is implemented.

By using the performance parameters determined during the initial sizing stage, following optimization problem is formulated for the mission and trajectory design and optimization.

The main title (on the first page) should begin 1-3/8 inches (3.49 cm) from the top edge of the page, centered, and in Times 14-point, boldface type. Capitalize the first letter of nouns, pronouns, verbs, adjectives, and adverbs; do not capitalize articles, coordinate conjunctions, or prepositions (unless the title begins with such a word). Leave two 12-point blank lines after the title. ∙ Maximize payload mass ∙ Subject to perigee altitude = 200 km

perigee velocity = 8,127 m/sec flight path angle at perigee = 0 deg

∙ Design variables: angle of attacks (14 design variables)

Mission design and trajectory optimization is performed to maximize the payload mass, while satisfying the given mission requirements. For the numerical optimization SQP of IMSL are employed. Figures 4 and 5 show the optimization history. Total

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rocket weight is 830.9 kg and the payload weight is 3.56 kg.

objective

-0.600

-0.550

-0.500

-0.450

-0.400

-0.350

-0.300

0 5 10 15 20 25 30 35 40

Iteration number

Objective

Figure 4. Objective function history

-0.150

-0.100

-0.050

0.000

0.050

0.100

0.150

0.200

0.250

0 5 10 15 20 25 30 35 40

x1 x2 x3 x4 x5 x6 x7

Figure 5. Design variable history

1st Stage Ignitiont=5sech=11,882mv=763.8m/sec

1st Stage Burnoutt=60sech=57,890mv=3528.9m/sec

2nd Stage Ignitiont=90sech=89,568mv=3433.7m/sec

Payload Fairing Seperation

t=155sech=149,822mv=5705.9m/sec

3rd Stage Ignitiont=214sech=191,421mv=5637.8m/sec

2nd Stage Burnoutt=124sech=121,731mv=5752.0m/sec

2nd Stage Seperaton

t=184sech=172,209mv=5669.3m/sec

3rd Stage Burnout & Orbital

Insertiont=241sech=200,138mv=8126.1m/sec

t=0sech=12,000mv=787.0m/sec

Figure 6. Optimized mission profile

3.3. Propulsion system Design 3.3.1. 1st Stage Hybrid Motor Design

The hybrid rocket engine, which uses a solid fuel and a liquid oxidizer, has in- remediate characteristics between the solid fuel and liquid propellant rocket engines. The popular combination of the fuel and the oxidizer is HTPB and LOX. Studies on the HTPB + H2O2, and PE/N2O + LOX combinations are going on. The concept of the hybrid propulsion system is described at Fig. 7. The liquid oxidizer is pumped into

the combustion chamber by the compressed gas. The valve located between the liquid oxidizer tank and the combustion chamber can supply, shut off and control the flow of the oxidizer, hence the ignition, extinguishments, and thrust control of the rocket can be possible. Specific impulse of the hybrid engines is 290 -350 sec, in-between the specific impulses of the solid and liquid propellant rockets[13].

Figure 7. Concept of the hybrid rocket propulsion system[14]

Based on the optimized trajectory of the rocket, first

stage hybrid rocket engine is designed [15]. Figure 8 shows the grain shape of the rocket fuel.

Figure 8. Grain geometry

Figure 9 shows the propulsion system configuration including the motor case and the liquid oxidizer tank.

The weight of the major subsystems obtained from the empirical equations, is summarized at Table 2.

Table 2. Mass of 1st stage main components

ComponentMotor case

Oxidizer tank

Nozzle & TVC

Structural supports

Mass(kg) 46.99 60.76 49.88 15.76

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Figure 9. 1st Stage hybrid propulsion system Second and Third Stage Solid Motor Design The grain shapes and components design of the second and third stages of the solid motor are shown at Figs. 10 and 11.

Figure 10. 2nd Stage solid motor

Figure 11. 3rd Stage solid motor 3.3.4. Wing Design

From the trajectory analysis, the maximum lift is estimated. The wing is design to produce the maximum required lift by considering the safety factor of 1.15. The design wing geometry is as follows, Wing area=3.81 m2; Root chord=2.47 m; Span=2.808 m.

Figure 12. Wing profile 3.3.5. Attitude Control System

The basic hardware components of the attitude control system of the space launcher are the inertial measuring unit (IMU), the on-board computer, and the driving actuator which can produce the required control moments.

In this study, the electro-magnetic actuator is used for the first stage fin control, and the thrust vector control is implemented for the second stage. For the third stage the thrust control using the spin motor is used. Figure 13 shows the schematic attitude control system of current study.

F-4E PhantomIMU

Flight Computer

-Navigation Calculation-Guidance Calculation

-Auto Pilot-Flight Sequencing

Air-Launching Rocket IMU

(LN-200S)

1st Stage Fin Actuator(EMA)

2nd Stage Thrust Vector Control

3rd Stage Spin Motor(Thruster)

Attitude, Velocity, Positon

DeflectionAngle

DeflectionAngle

Command

θ∆ V∆

Figure 13. Attitude control system

3.3.6. Digital Mockup Development

Digital mockup is fabricated using the CATIA V. 5.8. Figure 14 shows the part of the digital mockup of the rocket.

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Figure 14. Digital mockup using the CATIA 4. Conclusion and Future Works In this study, the preliminary design of the air-launching micro rocket was performed. Mission and trajectory optimization was performed and the propulsion system was designed based on the mission design. Total rocket weight was 830.95 kg, the payload weight was 3.56 kg. The designed propulsion system length is 5.89 m, hence the total rocket length will not exceed 7 m when considering the fairing length. Therefore, the designed rocket can be installed to the F-4E. Based on the current study, stability and control, and subsystem design will be performed. Moreover the system design by implementing the multidisciplinary design and optimization (MDO) techniques will be performed.

Acknowledgement This study was supported by fundamental research

progress (No. R01-2006-000-107400-0) of the KOSEF (Korea Science and Engineering Foundation) and BK21(Brain Korea) program.

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[3] http://www.ee.surrey.ac.uk/nano [4] Chul Park, "Prospects for Launch System Development in Korea," The First International Aerospace Technomart, Oct. 1996, pp. 79-104. [5] J-W. Lee, J. Y. Hwang, "The Conceptual Design of Air-Launching Micro Space Launcher, Mirinae-1," Journal of Korean Society for Aeronautical and Space Sciences, Vol. 29, No. 2, March, 2001. [6] http://www.astronautix.com/lvs/nots.htm [7] http://www.rocketry.com/mwade [8] Steven J. Isakowitz, Joseph P. Hopkins Jr., and Joshua B. Hopkins, International Reference Guide to Space Launch Systems, 3rd Ed. AIAA, 1999, pp.267-280. [9] Sobiezczanski-Sobieski, J., and Haftka, R. T., "Multidisciplinary Aerospace Design Optimization: Survey of Recent Developments ", AIAA-96-0711, 34th Aero. Sci. Meeting, Jan. 1996. [10] Jack D. Mattingly, William H. Heiser, Daniel H. Daley, Aircraft Engine Design, AIAA Education Series, 1987. [11] W. R. Roh, Y. D Kim, S. R. Lee H. J. Kim, "Trajectory Optimization of KOMPSAT Launch Vehicle Using Nonlinear Programming", Journal of Korean Society for Aeronautical and Space Sciences, Vol. 28, No. 1, Feb. 2001, pp.106-114. [12] Michael R. Mendenhall, Daniel J. Lesieutre, Steven C. Caruso, Marnix F. E. Dillenius, Gary D. Kuhn, "Aerodynamics Design of Pegasus: Concept to Flight with Computational Fluid Dynamics," Journal of Spacecraft and Rockets, Vol. 31, No. 6, November-December, 1994, pp.1007-1015. [13] W. J. Larson, Space Propulsion Analysis and Design, MacGraw-Hill, Inc., 1995. [14] http://www.isd.net/anowicki/SPBI101.HTM [15] S. T. Kwon, C. J. Lee, "Design of Air Launched Rocket System Using Hybrid Motor", Proceedings of the KSAS Spring Annual Meeting, Apr. 2002, pp 293-296.

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