Sounding Rocket

4th Responsive Space Conference RS4-2006-4004

Soundin oun ding g Rocket Rocket Technolo chn ology gy Demonstr mon stra ation tio n for fo r Small Small Satell Satellit ite e Launch Vehicl hi cle e Project ro ject John Tsohas, Lloyd J. Droppers, Stephen D. Heister Purdue University West Lafayette, IN

4th Resp Respon onsi sive ve Space Space Con Conference ference April 24–27, 2006 Los Angeles, CA

AIAA-RS4 2006- 4004

Sounding Rocket Technology Demonstration for Small Satellite Launch Vehicle Project John Tsohas1, Lloyd J. Droppers 2, and Stephen D. Heister 3 Purdue University, West Lafayette, IN, 47906

Purdue University is embarking on a program to demonstrate technologies critical to the development of a small satellite launch vehicle. The first phase of the program involves design, fabrication, testing and flight of a hybrid propulsion sounding rocket from the NASA Wallops flight facility. This paper details the design and test work that has been achieved to date. Propulsion work includes successful hot fire tests of a flight weight, 170 lbf thrust hydrogen peroxide / HTPB hybrid rocket motor at the Purdue rocket test facilities. The tests confirmed the structural integrity of the engine, verified the thermal insulation ablator design, helped determine solid grain regression rate and verified the engine performance characteristics with the internal ballistics simulation code. Detailed design of vehicle plumbing, structure, propulsion, avionics, and recovery subsystems has been completed. The rocket consists of a carbon-fiber composite aero-structure, welded aluminum oxidizer tank, and a fiberglass composite internal structure. A nitrogen blowdown system is used to provide the engine with oxidizer, and the recovery system has dual redundancy. In addition, detailed design has been completed on the ground support equipment used for remote loading and draining operations of liquid hydrogen peroxide to and from the vehicle, while monitoring critical vehicle parameters. Remote disconnect of umbilical cords, engine ignition, launch and aborts are also functions of the ground support equipment. A trajectory analysis and vehicle aerodynamics code was developed to design the vehicle geometry, stability, and mass allocation. Follow-on flights of the technology demonstration vehicle will include

the addition of a pressure fed cycle and a thrust vector system with associated guidance and control hardware and software. The second phase of the paper details the conceptual design of a small satellite launch vehicle designed to place 10 lb university or research payloads in low Earth orbit. In order to make use of the already existing rocket test facilities at Purdue and to keep test costs low, the thrust of the first stage engine was constrained to less than 10,000 lbf. To reduce costs associated with structural design, analysis and manufacturing, a three stage launch vehicle with a low propellant mass fraction for each stage (77%) would be designed. Hybrid propulsion would be used due to its relative simplicity and safety over liquid bi-propellant systems. Hydrogen peroxide would be used as an oxidizer due to the high density Isp and its non-toxic, and non-cryogenic properties. This would lead to a reduction in operations costs and increased safety in propellant handling in comparison with other candidate oxidizers. A small composite solid propellant third stage would provide the final delta-V at the desired orbital altitude. Thus, a three stage launch vehicle with a GLOW of 6,275 lb and 8,790 lbf thrust first stage engine would satisfy the above design requirements. Nomenclature

HTPB LOX c* ρ GOX

= hydroxyl terminated polybutadiene = liquid Oxygen = characteristic velocity = density = oxidizer flux

1

Graduate PhD Student, School of Aeronautics and Astronautics, Purdue University. Graduate Masters Student, School of Aeronautics and Astronautics, Purdue University. 3 Professor, School of Aeronautics and Astronautics, Purdue University. 2

1 of 15 4th Responsive Space Conference 2006

INTRODUCTION development program has been initiated at Purdue University School of Aeronautics and Astronautics to demonstrate technologies critical to the development of a small satellite launch vehicle. The first phase of the program involves the design, manufacturing and flight of a sounding rocket demonstrator at the NASA Wallops Flight Facility. The sounding rocket will serve as a test-bed for flight testing critical technologies including ground support, propulsion, structures, separation, recovery, telemetry, navigation, guidance and control subsystems. Each of these technologies will be demonstrated sequentially over a series of test flights. This will be achieved by increasing the test-bed subsystem complexity in each subsequent flight. Insequence flight testing will allow the designers to validate and fine-tune each of the aforementioned subsystems before adding more complexity, cost, risk and features to the technology demonstrator. By not having to fly all the vehicle subsystems on the inaugural flight, it is believed that a step-by-step flight validation approach will help control the design process, while concurrently reducing development cost and risk.

A

The second phase of the program involves design of a university built small satellite launch vehicle to place a 10 lb university or research payload in low Earth orbit. A similar bottom-top approach will be implemented in the design and flight validation of the launch vehicle by making use of technologies demonstrated during the first phase of the program. Consequently the third (final) stage of the vehicle will be validated in flight before the second stage is built. Following the same philosophy, the first stage will be built after successful flight validation of the third and second stages together. As mentioned earlier this incremental approach will serve to control the design, the cost and the risk associated with new launch vehicle development. This paper is divided in two parts. The first part of the paper details the design and test work that has been accomplished to date on the technology demonstration sounding rocket, while the second part details the conceptual design of the small satellite launch vehicle.

TECHNOLOGY DEMONSTRATOR ROCKET OVERVIEW

Hybrid propulsion was chosen over liquid and solid propulsion due to cost, complexity and reliability constraints placed early in the design process. Hybrid propulsion offers less complexity and higher reliability than liquid propulsion. Compared to solid propulsion, hybrid motors offer higher specific impulse and are safer in operation due to their ability to be shut down after ignition. Hydrogen peroxide was chosen as the oxidizer due to its high density Isp and its non-toxic, and non-cryogenic properties. This leads to safer propellant handling procedures which reduces operation costs compared to other oxidizers such as liquid oxygen. The performance of hydrogen peroxide outweighs the self-pressurization and relative safety of nitrous oxide as an oxidizer. In addition, Purdue University has the facilities as well as extensive experience with the use of hydrogen peroxide as a rocket oxidizer. HTPB is the fuel of choice due to its relatively high performance and regression rate, material properties and ease of manufacturing in comparison with other candidate hybrid fuels, as determined by thermo-chemical analysis. Table 1 compares performance of various hybrid oxidizer/fuel propellant combinations. The technology demonstrator vehicle is designed to reach an altitude of 25,000 ft, powered by a 250 lbf thrust engine, for thrust duration of 8 seconds. For the initial flights, the propellant feed system will operate in blow-down mode. More specifically, 1/4 of the oxidizer tank volume will be loaded with hydrogen peroxide, while the remaining 3/4 will be pressurized to a 600 psia MEOP with nitrogen. The pressure in the tank will decay as the liquid oxidizer exits the tank, thus leading to a gradual decrease in chamber pressure and consequently thrust. Follow on flights will include a pressurant tank and regulator in order to provide a constant 600 psia MEOP ullage pressure. The initial flights of the vehicle aim to verify the performance of the propulsion, structure, and recovery sub-systems, as well as ground support equipment for remote loading and draining of hydrogen peroxide oxidizer. The initial flights will not make use of active guidance, but instead will use fin stabilization. To maintain the stability margin, the vehicle will launch at an initial thrust-to-weight ratio of 4.5. Table A.1 in the appendix presents a detailed mass breakdown of the demonstrator vehicle. An overall vehicle schematic is presented if Fig. 1.


5,800 lbs was required to buckle a 0.056” thick quasiisotropic carbon-fiber tube. The oxidizer pressure vessel was designed with a safety factor of 2.0 X MEOP and was analyzed using hand calculations and ABAQUS finite element analysis software. The tank is manufactured with aluminum 6061-T6 body and aluminum 4043 weld filler material. The vessel has been proof tested to 1.5 X MEOP, and additional vibration tests are scheduled. A pressure relief valve set to 1.15 X MEOP will ensure that there is no over-pressurization of the pressure vessel. For added safety, a normally open solenoid vent valve is installed, to be closed during flight using lithium battery power. Fiber-glass is used to manufacture the nose cone, and composite honeycomb structure is used to make the guidance fins. The hybrid rocket engine is manufactured from aluminum 6061-T6, is designed with a factor of safety of 2, and has been proof tested to 1.5 X MEOP.

Figure 1. Schematic of sounding rocket demonstration vehicle

Table 1. Equilibrium composition Performance of Oxidizer/Fuel propellant combinations (220 psia Pc, optimally expanded for sea-level (sl) , or 3 expansion ratio 80 for ISP_vac, density in lb/ft at operating conditions.

Oxidizer/Fuel 98% H2O2 / HTPB 90% H2O2 / HTPB N2O / HTPB LOX / HTPB

84.9 82.9 50.2 67.7

O/F 6.1 6.7 7.6 2.1

Isp sl 228.4 222.5 219.5 244.1

Figure 2. Fiber-glass ring for mounting oxidizer tank to aero-structure.

Isp vac 331.0 325.5 323.7 366.1

STRUCTURAL DESIGN

The airframe of the demonstration vehicle consists of a 6” diameter carbon-fiber composite cylinder. The aluminum oxidizer tank is designed to fit inside the aero-structure, and is held in place by fiber-glass mounts which are bolted to the airframe as shown in Fig. 2. Stainless steel, ½” tubing connects the oxidizer tank to the hybrid rocket motor, which is fastened to the aero-structure by three 7075-T651 aluminum brackets, as shown in Fig. 3. A finite element analysis model of aluminum brackets is shown in Fig. 4. An additional fiber-grass mount is used to secure the hybrid motor in position. An eigenvalue buckling analysis shows that a force of

Figure 3. 7075-T651 aluminum brackets for mounting hybrid motor to aero-structure.


Sub-scale, 25 lbf Thrust Engine Design and Hot Fire Testing

Figure 4. Finite element analysis model of aluminum brackets.

ENGIRE DESIGN AND HOT FIRE TESTING

A hybrid rocket internal ballistics code was developed to model the performance of the motor, and to perform sizing of the rocket engine. The TEP (Thermal Equilibrium Program) was used to provide thermo-chemical data to the ballistics code. The thermo-equilibrium program receives inputs of the propellant formulation and characteristics, O/F ratio, area ratios and chamber pressure which it uses to calculate parameters such as c*, C F, Isp, Tc, and product molecular composition. The interior ballistics code uses a quasi-steady, lumped parameter approach assuming uniform fuel regression along the chamber length and well mixed flow at the throat. It assumes a spatially uniform chamber pressure and performs a mass flow balance with total nozzle flow rate and the injected oxidizer and vaporized fuel flow rates. The vaporized solid fuel flow and the total mass flow out are determined from c*, chamber pressure and throat area, with the oxidizer flow rate metered by a cavitating venturi. The regression rate is estimated based on a simplified model shown in Eq. 1 where a and n are empirically derived constants. From the geometry of the port, oxidizer mass flow rate, and regression rate constants, a chamber pressure and mixture ratio history are created which in turn yield the thrust profile.

r = aGox

n

(1)

A 25 lbf thrust, subscale engine was developed to acquire test data for characterization of the 90% H2O2/HTPB performance and regression rate under hot fire test conditions, and to validate the internal ballistics engine design code before attempting to scale up to the full-scale, flight-weight engine for the sounding rocket technology demonstrator. The thrust chamber was sized to produce 25 lbf of thrust at an MEOP of 440 psia. Table 2 presents relevant sizing parameters of the subscale, 25 lbf thrust engine. Because no data was found on the regression rates of H2O2/HTPB, the data from the LOX/HTPB propellant combination were used as a starting point for initial sizing. Fig. 5 presents a sub-scale engine schematic. Six hot fire tests were performed at four different initial Gox levels, varied from 0.2 lbm/(s-in 2) to 0.8 lbm/(s-in2) by changing the initial fuel grain port diameter. An average c* efficiency of 94% was obtained from the tests, indicating that acceptable levels of energy release occurred. Fig. 6 shows a picture of hot fire testing of the sub-scale engine. Table 2. Full-Scale Engine Parameters

Chamber Pressure Thrust Oxidizer Mass Flow Rate Initial O/F ratio Predicted average c* Fuel Grain Outer Port Diameter Chamber Length

440 25

[psi] [lbf]

0.1 5.5 5255

[lbm/s]

1.29 13

[in] [in]

[ft/s]

Figure 5: 25 lbf thrust sub-scale engine schematic


insulation are used to protect the injector face from combustion gases, as well as the walls of the post combustion chamber. In addition, high temperature RTV sealant is applied in the butt-seal joints to provide additional insulation and sealing.

Figure 6: Picture of 25 lbf thrust sub-scale engine hot fire test.

Full-scale, 170 lbf Thrust Flight-Weight Engine Design and Hot Fire Testing

The results from the 25 lbf thrust subscale engine tests were used to update the internal ballistics model which was used to design the full-scale engine. Flight vehicle performance requirements led to a 175 lbf thrust, flight-weight engine design, operating at an average O/F ratio of 5.25 and an oxidizer mass flow rate of 0.68 lbm/sec. Table 3 presents a summary of the design requirements for the full-scale engine. The flight-weight engine design differs significantly from the lab-weight, sub-scale ground test thruster. With the aid of ABAQUS finite element analysis software, the 170 lbf thrust flight engine structural design consists of an internal phenolic liner which acts as a bonding surface for the HTPB fuel grain, as well as an insulator between the combustion gases and the external aluminum chamber. As shown in Fig. 7, the phenolic tube is inserted inside a 6061-T6 aluminum chamber designed to contain an internal pressure of 440 psia MEOP. A flanged connection is used to bolt the injector piece to the aluminum chamber. To prevent leakage of combustion gases, a butt-seal between the injector piece and the phenolic liner is incorporated in the design. A secondary backup seal is created by placing a silicon o-ring between the injector piece and the combustion chamber as shown in Fig. 8. The nozzle uses ablative cooling, and is manufactured from pressed, discontinuous sheets of silica phenolic material. The seal on the aft end of the motor consists of a butt-seal between the nozzle and the phenolic liner (identical design to injector end seal). A secondary seal is created by a silicon o-ring placed between the nozzle and the aluminum chamber (as shown in Fig. 9). An aluminum retainer plate is used to hold the nozzle in place, and to provide positive pressure at the location of the buttseal joint. Sheets of carbon-filled EPDM ablative

The injector consists of stainless steel material, hollow cone spray, which was sized to provide 0.68 lbm/sec of oxidizer flow rate with a 20% pressure drop, and with the correct spray angle to impinge on the catalyst bed. Inserted at the top end of the fuel grain, a consumable catalyst bed (CCB) ignition system previously invented at Purdue causes hydrogen peroxide to decompose upon contact, increasing temperature and initiating combustion of the H2O2/HTPB propellant combination. Table 3. Full-Scale Engine Parameters

Chamber Pressure Thrust Oxidizer Mass Flow Rate Initial O/F ratio Predicted average c* Fuel Grain Outer Port Diameter Fuel Grain Length

440 175

[psi] [lbf]

0.68 5.25 4900

[lbm/s]

3.5 10

[in] [in]

[ft/s]

A series of hot fire tests were performed in order to verify the structural integrity of the new engine, to verify thermal integrity of injector face, postcombustion chamber and nozzle, as well as to ensure that no leakage was taking place past the primary and secondary seals. The tests also aimed to obtain regression rate and combustion performance data, and to verify/calibrate the hybrid rocket motor internal ballistics code. A total of three hot fire tests successfully took place, for burn durations of 3.0, 3.5 and 6.5 seconds. No problems were encountered with the structure, and upon close examination of the butt and o-ring seals, no sign of leaks, material degradation or charring was observed. There was minimal ablation of the carbon-filled EPDM insulation on the injector face. However, on the third hot fire test, after a cumulative time of 13 seconds, chipping of the discontinuous silica phenolic sheets was observed at the nozzle throat. Fig. 10 shows a picture of the full-scale engine hot-fire testing. The thrust and chamber pressure measurements for the three hot fire tests (shown in Fig. A.1 of the appendix) were in close agreement to the predicted values. The actual regression rates were higher than


predicted, showing a 63% increase over the subscale engine tests. The higher than expected regression rates are attributed to air bubbles formed during casting of the HTPB fuel grain, which cause uneven burning and an increase in the burn surface area. For this series of tests, the average specific impulse was 218 seconds (moderately over-expanded sea-level thrust).

Injector Assembly

Combustion Chamber Top Cap

Figure 8. Injector end seal and assembly detail.

Consumable Catalyst Bed

Aluminum Motor Casing

HTPB Fuel Grain

Figure 9. Nozzle end seal and assembly detail. Aft Mixing Section Paper Phenolic liner Silica Phenolic Nozzle Viton O-ring Nozzle Retaining Ring

Figure 7. Full-scale, 170 lbf thrust, flight-weight engine assembly drawing.

Figure 10. Picture of flight-weight, 170 lbf thrust engine hot fire test.


RECOVERY SUB-SYSTEM

The incremental development of the technology demonstration sounding rocket relies heavily on multiple launch and successful recovery of the flight hardware. This calls for the development of a very robust and reliable parachute recovery sub-system. Following a recovery system trade study, the final design calls for a dual deployment system, with drogue parachute deployment at apogee, and primary parachute deployment at 1,200 ft altitude. System redundancy is achieved by making use of two completely independent recovery modules for parachute ejection. Each module contains a lithiumion battery, an R-DAS flight computer and four pyrotechnic ejection charges (2 per parachute). The recovery sub-system will be tested on a separate flight prior to being used on the demonstrator vehicle. Fig. 11 shows a layout of the recovery sub-system.

Figure 11. Recovery sub-system layout.

temperature are constantly monitored to verify that the hydrogen peroxide is not undergoing unexpected decomposition. Launch is initiated by opening an on board electrically actuated solenoid valve which allows hydrogen peroxide to flow into the main engine where ignition takes place upon contact with the consumable catalyst bed (CCB). In the event of an abort, the GSE has the capability of remotely draining the hydrogen peroxide from the launch vehicle into a specially designed dump tank on the ground, by simply shutting off the pressurization source, and opening a dump valve. In the event where an abort is called after the quick disconnect has been disengaged, depressurization of the vehicle occurs by opening a remotely actuated solenoid vent valve on the tank. To ensure safety in launch operations, all circuits of the GSE and launch vehicle are designed to be fail-safe. In the event of an unexpected power outage, all solenoid valves return to their normal positions (normally open or closed) to allow venting of the tanks and automatic draining of the oxidizer from the launch vehicle directly into the dump tank. Fig. A.1 in the appendix shows a plumbing and instrumentation diagram for the GSE system. Valve control and monitoring of pressures and temperatures on the GSE/launch vehicle system is controlled by Labview software.

AVIONICS, GUIDANCE, CONTROL

Guidance and control will be implemented on followon flights of the technology demonstration vehicle, following flight verification of the recovery, propulsion and structures sub-systems. A liquid injection thrust vector control system (LITVC) with associated guidance algorithms, software and hardware is currently being developed. GROUND SUPPORT EQUIPMENT

The ground support equipment (GSE) is used for remote loading and draining of hydrogen peroxide to and from the launch vehicle. The system consists of 5, ½” solenoid valves, 2 pressure regulators, 4 pressure transducers, 2 thermocouples, 4 check valves and associated ½” plumbing lines. Regulated nitrogen pressure is used to pressurize a hydrogen peroxide storage tank to the desired ullage pressure of 600 psia MEOP in order to feed liquid oxidizer through a series of solenoid valves into the launch vehicle oxidizer tank. The nitrogen gas continues to flow until all pressures in the system reach equilibrium. The oxidizer fill line is remotely disengaged from the vehicle by a remotely actuated quick-disconnect valve. To ensure safe launch operations, the oxidizer tank pressure and

OVERVIEW OF CONCEPTUAL SMALL LAUNCH VEHICLE DESIGN

The second phase of the program involves the design and flight of a university built small satellite launch vehicle capable of delivering a 10 lb mass payload to Low Earth Orbit. Technologies developed and proven in flight under the first phase of the program will be implemented in the design of the orbital vehicle. Performance data from flights of the demonstrator vehicle will provide baseline numbers from which to begin the design. The following pages present a conceptual design based on test data obtained to date.


DESIGN PHILOSOPHY

The design drivers for launch vehicle development depend on the needs and resources of the particular organization. Profit is a typical design driver for commercial space organizations, research and development is a typical design driver for government organizations, and defense and security applications are for military organizations. These design drivers are prevalent and visible in all aspects of the rocket and launch infrastructure design, and the same is true for a university organization. The Design Philosophy of a university built launch vehicle is mainly influenced by the needs, resources and capabilities of the university organization, its students and its faculty. The primary driver for a university is to provide its student body with the education, and this project will provide practical engineering experience to design, build, test and fly a launch vehicle into Earth orbit. Through this design process, university students and faculty will obtain the opportunity to conduct research, and to test innovative design concepts in multiple sub-orbital and orbital flights. These design drivers lead to the necessity of orienting the program for multiple design and flight test cycles, which will lead to increased student exposure. In addition to student education and exposure, funding for the effort will most likely come from corporate sponsors which are inclined to invest in a program with clear, high visibility milestones, such as those provided by test launches. Moreover, the constraint of working with relatively small groups of students inhibits the ability of a parallel, rather than series style development program, due to a much larger number of students required to complete a parallel vehicle design effort. Due to these factors an in-series vehicle development program is proposed. Using this philosophy, the vehicle will be designed in three separate stages, each capable of being manufactured and tested individually. The 3 rd stage of a 3-stage vehicle will be designed, built and tested first. Upon flight verification of the 3 rd stage, the 2 nd stage of the launch vehicle will be built and assembled. In the same fashion, following successful flight verification of stage 2, the 1 st stage will be manufactured, assembled and flown. Serial manufacturing, integration, testing and launch of each stage individually will help work out the kinks in the sub-systems without putting the fully integrated launch vehicle at risk. Moreover, smaller scale projects within the framework of a larger program will allow students to participate in a project

from its inception to its completion, while still being students. The main challenge with a serial development approach is that of maintaining design heritage and institutional knowledge. When one generation (class) of designers has graduated the next needs to be able to pick up where the previous one left off, which requires a dedicated management team and extensive documentation. The conceptual design calls for 98% H 2O2/HTPB hybrid propulsion and LITVC vector control on the 1st and 2nd stages. The 3rd stage uses solid propulsion with spin-stabilization to boost the payload into a low Earth orbit. As mentioned earlier, the plan is to test each of the individual stages as ground launched suborbital rockets. The 1 st stage will not require much modification for sub-orbital testing since it is designed with active guidance, is mostly re-usable and incorporates a parachute recovery sub-system. For sub-orbital testing, the 2 nd and 3rd stages will need to be modified with the addition of a parachute recovery module and a nozzle designed for sea-level operation. Due to its lack of active guidance, the 3 rd stage would require a thrust-to-weight ratio of 4 to be launched from the ground using temporary, additional fin stabilization. Therefore, the 3 rd stage will have a higher thrust-to-weight ratio than what is typically for the final stage of a launch vehicle. In addition, the test launch of the individual stages will provide available payload mass and microgravity time for universities to conduct student built microgravity experiments. A low propellant mass fraction is a major constraint on a university built launch vehicle. When based solely on performance a high propellant mass fraction is clearly preferable, however when one considers the primary mission drivers, such as those of educating students through practical design experience, it becomes clear that high propellant mass fractions are not feasible. The relative inexperience and limited time commitment of student designers, leads to the necessity of adding higher than normal factors of safety. The manufacturability of parts by in-house or relatively inexpensive local machine shops as compared to aerospace grade certifiable machine shops leads to higher factors of safety as well. Furthermore, less infrastructure is required to build a launch vehicle with a relatively higher safety factor, and a lower propellant mass fraction. Design with lower propellant mass fractions will lead to a more accessible design, lower development cost, at the expense of vehicle performance (lower payload mass to orbit).


CODE DEVOLOPMENT

Computer code was developed to aid in the conceptual design of a small satellite launch vehicle. The basic scheme involves solving the ideal rocket equation for a given mission ∆V, assuming a propellant mass fraction. The first part of the simulation runs an engine internal ballistics code in conjunction with a vehicle trajectory code to estimate ∆V losses due to gravity and drag, subsequently updating the required mission ∆V. The inert masses are tabulated using both calculated and historical mass estimates, from design handbooks and experience from the demonstrator vehicle, subsequently updating the assumed propellant mass fraction. The vehicle simulation is iterated until it converges on a design solution. IDEAL ROCKET PERFORMANCE

The premise of this code is the ideal rocket equation in the following form. ⎛ m ⎞ ∆V = − Isp * g ln⎜⎜ f ⎟⎟ ⎝ mi ⎠ MR =

m prop

the dynamic pressure calculated from density input from the Standard Atmosphere as well as velocity and drag coefficient (C d ) values, based on a subsonic compressible aerodynamics model. The model is than modified for transonic flow using Prandtl-Glauert, and as a conservative assumption, is assumed to be 0.4 for supersonic conditions. The gravity force is assumed to be constant (32.2 ft/s2) throughout the entire flight. The code allows for thrust vectoring at any given angle, but for this analysis the engine is simply commanded to fly at a constant flight path angle, and then to perform a gravity turn. The code uses a finite difference model whereby the position, velocity, and acceleration of the vehicle are calculated based upon input of force balance, position, and velocity from the previous time step. The code is run past engine burnout, allowing the vehicle to coast to apogee. Based upon the altitude and velocity achieved at apogee, a ∆V required to circularize the orbit is calculated, which is then iterated in conjunction with the ideal rocket performance code to converge upon a design.

m f mi

= f prop * m pl *

( MR − 1) 1 − MR * (1 − f prop )

Where, MR is mass ratio, m f is final mass, m i is initial mass, and f prop is propellant mass fraction. The code receives input of payload mass (m pl), inert mass fractions and total ∆V, and outputs propellant mass, inert masses and ∆V for each stage. Initially the upper stage values are calculated, and the total upper stage is used as the payload mass for the lower stage, and similarly the 2 nd and 3rd stages total mass is used as the first stage payload mass. For vehicle optimization the code runs a wide range of cases and finds the minimum mass solution for the ∆V breakdown.

TRAJECTORY

The trajectory code calculates the vehicle trajectory based on engine performance input from the internal ballistic code and with estimates of vehicle mass and geometry from the ideal rocket performance code. An equations of motion force balance is performed which calculates acceleration using a two dimensional flat Earth model and on gravity, thrust and drag models. Thrust is calculated using output from the internal ballistics code, and the ambient pressure at each altitude using the 1962 U.S. Standard Atmosphere model. Drag is calculated using

INERT MASS

Using historical trends, an inert mass break down for the two hybrid rocket stages is calculated and compared to the allocated inert mass for each stage. The inert mass code uses historical values based on SPAD as well as development data from the demonstrator vehicle (phase 1). The inert masses are broken down into 5 sub-sections, including propellant tank mass, ullage mass, pressurant mass, structural mass, and extraneous mass such as valves, plumbing and wiring. The propellant tank mass is calculated using the pressure vessel performance-factor approach, where the tank mass factor φtank is taken from historical values. The tank mass factor φtank is shown in Eq, 2, where V tank is tank volume, and P b is burst pressure. The helium pressurant mass is calculated based on isentropic expansion of the pressurent into the oxidizer tank. Pressurant tank mass is calculated in a manner similar to the propellant tank mass. Based on historical data, 10% of the inert mass was assumed for structural mounts, airframe, bolts etc. In addition, valves and other components were estimated as a further 10-15% of the inert mass. Using the calculated inert mass, the propellant mass fraction can be updated for further iteration with the ideal rocket performance code.


mtank =

pbV tank

φ tank g 0

(2)

LAUNCH VEHICLE MASS BREAKDOWN

The launch vehicle mass breakdown was calculated using the codes and procedures described in the previous section. The design was constrained to placing a 10 lb payload at a minimum altitude of 93 miles, assuming a maximum 1 st stage thrust of 10,000 lbf, based on the testing capabilities of the High Pressure Laboratory at Purdue University. Given the above constraints, as well as Isp and engine performance data from preliminary phase 1 hot fire testing, the code found the minimum propellant mass fraction that would achieve mission objectives. A minimum propellant mass fraction of 77% was converged upon. Table 4 presents the initial vehicle mass breakdown for each stage. The vehicle gross lift-off weight (GLOW) is equal to 6,275 lb, making use of 680 lb of HTPB fuel, and 4,082 lb (350 gallons) of 98% concentration hydrogen peroxide oxidizer. The total length of the vehicle equals 27.7 ft with a base diameter of 3.0 ft. Fig. 12 shows the calculated dimensions for each stage, and approx. size of major components. The propellant mass fraction will be 77% which is considerably lower than the 85% (approximate average) mass fraction of modern day launch vehicles. As mentioned earlier, a low propellant mass fraction is desirable due to design limitations and construction techniques available to a university organization. The composite oxidizer and pressurant tank masses, thrust vector control hardware mass, and avionics system masses were estimated using historical trends. rd

3 STAGE

The serial design philosophy states that our first vehicle design will be the 3 rd stage. The third stage has a low total mass and for this reason, solid propulsion will be able to achieve the inert mass fraction limitation due to its lack of valves, plumbing and other components. Based on historical trends, an Isp of 290 sec was assumed for the third stage solid motor, using an nozzle with 100:1 expansion ratio. The third stage motor will be ignited at the desired orbital altitude, providing the delta-V required to achieve orbital velocity. Four, cold gas roll thrusters on the 2nd stage spin the vehicle following 2 nd stage burn-out. Following stage separation, the 3rd stage is

spin-stabilized and thus requires no active guidance. A propellant mass fraction of 0.77 was assumed for the 3rd stage. An inert mass of 18.7 lb includes the SRM casing, avionics, payload mount and the satellite release mechanism. The solid propellant rocket motor is designed to produce an average thrust of 370 lbf, for total burn duration of 49 seconds.

nd

2 STAGE

A 98% H2O2/HTPB hybrid engine is used as primary propulsion for the 2nd stage of the launch vehicle. Data from the demonstrator vehicle engine was used as an initial design point, with a reduced chamber pressure of 200 psi and increased hydrogen peroxide concentration from 90% to 98%. An Isp v of 320 sec is estimated from the NASA thermal equilibrium program, and a 95% c* efficiency based on sub-scale and full-scale demonstrator vehicle engine tests. The hybrid engine is designed to produce an average thrust of 1,230 lbf, for total burn duration of 157 seconds. The second stage has a total mass of 879 lb, with 605.9 lb of propellant and 181 lb of inert mass. Table 5 shows a further inert mass breakdown of the 2nd stage. The mass of the pressurant and oxidizer tank were both estimated by assuming a conservative pressure vessel performance factor of 21,300 ft, only moderately better than large DOT composite overwrapped pressure vessels. The structural performance of the pressure vessel can be improved given the availability of resources and appropriate corporate sponsorship. The engine mass was estimated from the demonstrator vehicle engine thrust-to-weight ratio of 31.3. This value can be improved given the larger size of the vehicle. The propellant ullage mass was assumed to equal 2.5% of the stage propellant mass based on Huzel and Huang design guidelines. Miscellaneous mass include valves, plumbing, thrust vector control, wiring, etc. is assumed to equal 15% of the inert mass. To allow for unforeseen increase in the mass of the 2 nd stage, a weight growth margin of 53.8 lb was allocated.

Table 4. Overall Vehicle Mass Break Down, lbs

Mtot stg1 Mprop stg1 5397 86% GLOW

4156 77% STG1

Mi stg1 Mtot stg2 Mprop stg2

Mi stg2 Mtot stg3 Mprop stg3

Mi stg3 Mpayload M GLOW

1241 786.9 23% 12.5% STG1 GLOW

180.7 81.2 23% 1.3% STG2 GLOW

18.7 10.0 6275.1 23% 0.16% STG3 GLOW

605.9 77% STG2

62.5 77% STG3


Table 5. Stage 2 inert mass breakdown, lbs. Ullage Mass, lbm 15.1 8.4% Oxidizer Tank mass, lbm 16.2 9% Engine Mass, lbm 28.1 15.6% Pressurant Mass, lbm 5.5 3% Pressurant tank, lbm 16.8 9.3% Structural Mass, lbm 18.1 10% Miscellaneous mass, lbm 27.1 15% Mass Growth, lbm 53.8 29.7% Stage 2 Total Mass, lbm 180.7 100%

st

1 STAGE

The 1st stage design is similar to that of the 2 nd stage, but on a larger scale. The H 2O2/HTPB hybrid engine will operate at a chamber pressure of 600 psia to improve specific impulse to an average of 260 sec over the entire flight envelope of the 1 st stage. The engine is designed to produce 8,790 lbf of thrust for total burn duration of 123 seconds, well within the capabilities of the rocket test facilities at Purdue University. Miscellaneous mass accounts for 10% of the inert mass due to the larger size of the 1 st stage, which effectively reduces the fraction of vehicle mass allotted to valves and other small components. The helium pressurant tank is pressurized to 9,000 psia for the regulated pressure fed system. Table 6 presents the inert mass breakdown of the 1 st stage. Fig. A.3 in the appendix shows plots of altitude, velocity and thrust versus time for the 1 st stage. The use of a proprietary altitude compensation nozzle is being considered in the conceptual design trade studies in order to boost specific impulse during ascent through the atmosphere. Finally, to further increase Isp, alternative fuels to HTPB are being investigated such as DCPD. Table 6: Stage 1 inert mass breakdown, lbs. Ullage Mass, lbm 103.9 8.37% Oxidizer Tank mass, lbm 201.6 16.2% Engine Mass, lbm 281.1 22.6% Pressurant Mass, lbm 69.7 5.6% Pressurant tank, lbm 230.7 18.6% Structural Mass, lbm 124.1 10% Miscellaneous mass, lbm 124.1 10% Mass Growth, lbm 105.9 8.5% Stage 3Total Mass, lbm 1241 100%

Figure 12. Overall layout of small satellite launch vehicle.


CONCLUSIONS The design drivers for launch vehicle development depend on the needs and resources of the particular organization. The primary driver for a university is to provide its student body with the education, and the practical engineering experience to design, build, test and fly a launch vehicle into low Earth orbit. Through this design process, university students and faculty will obtain the opportunity to conduct research, and to test innovative design concepts in multiple sub-orbital and orbital flights. A sounding rocket demonstrator is being designed and tested at Purdue, to serve as a test-bed for flight testing technologies critical to the development of a small satellite launch vehicle. The complexity of the subsystems launched will increase with each subsequent flight. In-sequence testing will allow the designers to validate and fine-tune the subsystems before adding more features, cost and complexity to the demonstrator vehicle. By not having to fly all the vehicle subsystems on the inaugural flight, it is believed that a step-by-step flight validation approach will help control the design process, while concurrently reducing development cost and risk.

The second phase of the program involves the design of an orbital vehicle by making use of technologies demonstrated during the first phase of the program. A “serial” development approach will be implemented whereby the third (final) stage of the vehicle will be validated in flight before the second stage is built. Consequently the 1st stage will be built after successful flight validation of the third and second stages. As mentioned earlier this incremental approach will serve to control the design, cost and risk associated with new launch vehicle development. To reduce costs associated with structural design, analysis and manufacturing, a three stage launch vehicle with a low propellant mass fraction (77%) would be designed and manufactured. Hybrid propulsion would be used due to its relative simplicity over liquid-bi-propellant systems. In addition, the 98% H 2O2/HTPB propellant combination offers high density Isp as well as nontoxic and non-cryogenic properties which leads to increased safety and a reduction in operation costs. A small composite solid propellant third stage would provide the final delta-V at the desired orbital altitude of 93 miles. Finally, a three stage launch vehicle with a GLOW of 6,275 lb powered by an 8,790 lbf thrust first stage engine would satisfy the mission design requirements.

ACKNOWLEDGEMENTS The authors would like to thank the members of the Purdue University Hybrid Sounding Rocket team, especially Jeremy Corpening and Michael Grant, as well as Scott Meyer for their help and support. We would also like to acknowledge the sponsors to the Hybrid Rocket project and thank them for their support: ATK Thiokol, SpaceX, Aerojet, Purdue School of Aeronautics and Astronautics, and Rocky Mountain Wireline Service.

REFERENCES 1

Gordon, S., McBride, B, Computer Program for Calculation of Complex Chemical Equilibrium Compositions, Rocket Performance, Incident and Reflected Shocks, and Chapman-Jouguet Detonations, NASA SP273, 1971. 2

Ben-Yakar, Adela; Gany, Alon. “Hybrid Engine Design and Analysis.” Israel Institute of Technology; Haifa, Israel. AIAA 1993. 3

Wernimont E. J., Heister S. D., “Combustion Experiments in Hydrogen Peroxide/Polyethylene Hybrid Rocket with Catalytic Ignition,” Journal of Propulsion and Power , Vol.16, No. 2, 2000, pp 318-326. 4

Sutton, G., Biblarz, O., Rocket Propulsion Elements and Design, 7th ed, John Wiley & Sons, New York, 2001, pp 64. 5

Wernimont, E. J., Meyer, S. E., and Ventura, M. C., “A Hybrid Motor System with a Consumable Catalytic Bed, A Composition of the Catalytic Bed and A Method of Use,” U.S. Patent Application 08/623,937,Filed March1996. 6

Humble, R., Henry, G., Larson, W., Space Propulsion Analysis and Design, McGraw-Hill, New York, 1995, pg 372. 7

Estey, P.N, Whittinghill, G.R., “Hybrid rocket motor propellant selection alternatives,” 28 th Joint Propulsion Conference and Exhibit , 1992, AIAA-1992-3592. 8

Larson, W., Wetrz, J, Space Mission Analysis and Design, Microcosm Press, 1999 9

Huzel, D., Huang, D., Modern Engineering for Design of Liquid-Propellant Rocket Engines, AIAA, 1992


APPENDIX

Table A.1. Sounding rocket mass breakdown, lbs

Recovery Sub-system (lb): drogue chute primary parachute primary/drogue shock cords drogue & main plunger fasteners - recovery system primary/drogue pyro squibs flight computer #1 and #2 additional electronics lithium ion battery #1 and #2 electrical wiring avionics housing structure Recovery Sub-syst em Total: Structure Sub-system (lb): nose cone fuselage aero-structure Fins oxidizer tank pressure relief valve contingency vent valve pressure transducer temperature thermocouple ox tank fittings (vent section) quick disconnect and fittings fasteners - ox tank/fuselage TCA mount Bulkhead recovery module aero-structure Lithium battery pack Struct ure Sub-system Total: Propulsio n Sub-system (lb): main valve thrust chamber feed lines thrust chamber assembly fasteners - thrust chamber Propulsio n Sub-system Total: VEHICLE DRY MASS (lb ): OXIDIZER MASS (lb): VEHICLE GLOW (lb):

0.58 0.61 1.0 0.4 0.6 0.2 0.1 0.2 0.6 0.2 0.7 5.2 1.5 6.0 0.39 10.75 0.4 0.38 0.35 0.1 0.3 0.6 0.1 0.15 0.15 3.5 0.5 25.2 2.0 0.3 11.4 0.1 13.8 44.2 11.5 55.7


Figure A.1. Thrust and chamber pressure measurements from flight-weight, 170 lbf thrust engine hot fire tests.

Figure A.2. Ground support equipment plumbing and instrumentation diagram.


5

3

a) Altitude vs. Time

x 10

] t f [ 2 e d u t i t 1 l A

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] s / t f [ 4000 y t i c o 2000 l e V

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Figure A.3. Altitude, velocity and thrust versus time for small satellite vehicle 1 stage.


Sounding Rocket

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