The idea of designing a hybrid sounding rocket to carry a small component of payload into microgravity is indeed a luminous resolution. The design would inherently show the concept of real world situation thus representing the requirements. Here, in this project, we design a hybrid sounding rocket that will dependably meet the requirements to perform its mission. This involves traditional aerospace engineering field such as structures, propulsion, flight mechanics, orbital mechanics, avionics, attitude control, sensors and system design optimization. Besides, economics and plan will also play a key role in determining design feasibility.
Sounding rockets which are also known as research rocket is designed to carry a payload, take measurements and carry out scientific experiments. The rockets are designed to carry instruments to typical operational altitudes which are in between 50 to 1500km above the surface of Earth. The advantages of sounding rockets in some research area are their low cost, short lead time as well as the ability to conduct study in areas inaccessible to either satellites or balloons.
A hybrid rocket has a rocket engine which uses two different matter of propellant to perform. One is solid and the other is either gas or liquid. The concept of a hybrid rocket consists of a pressure tank containing the liquid propellant, a combustion chamber containing the solid propellant and a valve separating the both. Commonly, the liquid propellant is oxidizer and solid propellant is the fuel as solid oxidizers are low in performance compare to liquid oxidizers. There are more advantages in the use of hybrid rocket compare to liquid and solid rockets. Technically, hybrid rockets are mechanically simpler, have denser fuels and are metal additives which are beneficial in increasing the specific impulse. Moreover, hybrid rockets are more controllable, safe as in using non-toxic oxidizer and contribute less in explosion hazard.
800px-Hybrids_big.png
Figure 1.1 - Concept of Hybrid Rocket Propulsion System
www.spg-corp.com/space-propulsion-group-resou...
PROJECT REQUIREMENTS
The aim of this project is to design a hybrid sounding rocket which would meet the following requirements:-
Mission Requirements
Rocket must be capable of carrying a 20kg payload to an altitude of 150km.
The payload should be in a useable microgravity environment for as long as possible.
Gravity levels reduced to 10-2 or less are desirable.
Absence of angular body rate is mandatory for generation of microgravity environment.
Provide a layout for the payload module that is convenient for assembly and disassembly, servicing, calibration and testing.
Retrieval of results and data collected by the microgravity experiment equipment.
Safe recovery of the payload, i.e. the microgravity experiment equipment.
Product Assurance Requirements
Reliability of the proposed designs or chosen techniques.
Consider availability of products proposed to be brought into use.
Maintainability, calibration and testing of subsystem components.
Performance Requirements
Ranging, pointing and measurement accuracy.
Stability
Data storage and transmission capacity and on-board processing.
Effective and timely functioning of the separation mechanism, de-spin mechanism and parachute deployment.
Reliable tracking and recovery system.
Safe and efficient hybrid rocket propellants and a dependable motor design.
Structural strength and stiffness.
Physical Requirements
Payload, launch vehicle and propellant masses.
Launch Vehicle dimensions.
Power/ thrust required to meet the 150km range requirement.
Internal torques and disturbances.
Cost Effective Services for the customer.
DESIGN CONFIGURATION
Rockets are made up of stages which are multistage and single stage. Staging is being a major contribution to the rocket design to enhance the overall performance whilst still consuming smaller motors. This is performed by emitting unused mass along the escape as the propellant is spent. The design for the hybrid sounding rocket in this project is a single stage rocket. In order to carry a small payload according to the requirement, a single stage rocket is an appropriate shoot up compared to a multistage. However, there are pros and cons comparing to a multistage design. A single stage would have a lower operating cost, better reliability, straightforward separation and enhanced safety. In addition, as this stage is would be recoverable thus all the factors of having this stage would be favourable. Besides that, by having this single stage system, many complications to the design could be avoided such as separating costly engines and structure, complexity of reinstating a large stage and higher improvement cost of two separate vehicles.
A weight decision matrix is done in Table …. Below to decide between a single stage design and multistage design.
Specification
Points
Single stage rocket
Multistage rocket
Cost
9
7
63
3
27
Structure
7
8
56
6
42
Weight
6
4
24
7
42
Construction
3
6
18
4
12
Sub element
1
6
6
3
3
Efficiency
5
4
20
6
30
Total
187
156
Table
Cost and time constraints are the most essential part of the project which has to be given top priority. This constraint play major role in a decision as the other listed specification directly related to cost. Structural complexity and sub element contribute to the total weight and construction difficulties. Even though efficiency of the rocket is vital, trade off with cost and time has rated efficiency constraint low.
Single stage rocket mostly suite low altitude research as it provides cost and time efficiency in term of construction as well as weight constraint. Meanwhile, multistage rocket for this project would not be suitable as it add more complexity and time consuming in term of construction.
SEPARATION MECHANISM
A weight decision matrix is shown to decide on the type of mechanism:
Specification
Points
Mechanical system
Pyro-technique devices
Joint rotation
7
7
49
5
35
Simultaneity
8
3
24
6
42
Reliability
8
7
56
6
42
Low shock levels
10
8
80
4
40
Weight
10
7
70
7
70
Cost
9
6
59
5
45
Total
338
274
As the separation mechanism act as a medium that deliver the payload to its final altitude, it's important to make sure no damage is caused to payload during the separation process. Therefore, low shock level and extra weight the mechanism adding to the rocket is essential trade off. Timing and accuracy of the separation mechanism is essential to deliver successful project. Besides that, non spinning payload study is the requirement of the projects; joint rotation of the mechanism has to be given more priority.
Mechanical system has less commitment toward adding complexity in term of construction and avionics requirement. Even though its reliability and joint rotation criteria are close to the pyro-technique, the amount of time and cost that will be spend on a low altitude research rocket will be less as the mechanical system is applied to the sounding rocket.
Separation mechanism is an instrument which involves in releasing stages in a rocket. In designing the separation mechanism, some significant factors need to be measured: Sufficient clearance between the separating bodies; No damaging shock loads provoked in the structure and payload; No contamination such as extreme or destructive debris from the operation. There are three basic elements to be considered for the operation of the separation. An actuator which is either electric or pyro based is the first element to trigger the event. The second element is release mechanism for physical separation. Jettisoning system is part of the separation element used to provide the essential relative separation velocity. There are two commonly used rocket separation mechanism which are the mechanical system and pyro-technique devices. The mechanical separation mechanism works internally without internal radiated energy. Meanwhile, the pyro system works in the collaboration of firing electrical charges. In the case of this project requirement, ball and lock system which is the mechanical separation mechanism is the most suitable implement to be used.
Ball and lock system consist of three main components: The upper and lower rings which are adoptable and attached together by steel; A retainer ring which has escape holes for balls; Balls which are hold by a retainer ring. The holes in the retainer ring are set to a balance angular in the locked condition. The pyro-thruster swivel the retainer ring which causes it to be imbalance during release. The system has a stopper to limit the rotation of the retainer ring. An electric command from the central sequencer of the rocket instigates the pyro thrusters. Jettisoning system used is a helical compression spring in between the flanges to report the vital differential separation velocity. A clean separation is achievable in this system as the lower stage outer ring is endowed with holes all the way through for the balls in the locked condition. Meanwhile, the upper stages adapter ring comes with a conical ball seat. The radial element of the spring forces eventually pushes the balls away from the ring and discharges the inner ring which performs a clean separation.
The advantages of this mechanism comparing to many others are the good joint stiffness, lightweight construction, tuneable jettisoning velocity, debris free actuation and redundancy in launch to increase the reliability of the system. Moreover, ball and lock system generates low release shock. However, a good understanding of system dynamics would be essential in order to identify the contribution of shock from different sources.
Figure : Ball and Lock System
http://www.raes.org.uk/pdfs/3007.pdf
ATTITUDE STABILISATION WITH SPIN
The attitude of a body can be stabilised by simple passive methods such as imparting spin to it. A rigid body with its angular velocity parallel to its major axis (the principal axis passing through the centre of mass, having max. moment of inertia) will maintain this axis in a fixed direction with respect to inertial space in the absence of external torques. [2] Hence Spin Stabilisation is considered a useable attitude control technique for our rocket, which would correct instabilities by creating angular momentum.
Various means of imparting spin could be:-
Canting the fins, i.e. giving the fins a slanting edge.
Air-foiling all the fins, i.e. all the fins facing the same direction.
Releasing the exhaust gases at an angle (this technique was used in earlier times).
However, the payload requires a microgravity environment for the research experiments, which implies the following two conditions must be met:-
Absence of net acceleration force (force gradient) acting on the body.
Absence of angular body rate (spin).
The net acceleration force acting on the body is significantly reduced in the free-fall state above Earth's sensible atmosphere. And to efficiently utilise the time above the Earth's atmosphere, a positive means of controlling angular body rates must be provided for the payload. [3] Hence the payload must be de-spun to sustain the microgravity requirements. Various active (with sensors) and passive (without any sensors) Attitude Control Systems (ACS, were considered to support this requirement, such as:
Three Axis Rate Control System and Guidance Control System
Yo-Yo De-Spin Mechanism
De-spun Platform
The passive or semi-passive attitude control systems (particularly the Yo-Yo De-spin Mechanism) were found to be lighter, simpler and more cost-effective solutions to meet the customer requirements. Since the de-spun platforms require servo motors for stabilisation, the yo-yo mechanism is considered a convenient option for this project. The Weighted Decision Matrix depicted below, was created for selection of the de-spin mechanism.
Mechanism
Features
Weight
Complexity
Cost
Effectiveness
Rate Control System
4
4
3
8
De-spun Platforms
7
8
6
7
Yo-Yo De-spin Mechanism
8
9
8
7
Yo-Yo- De-Spin Mechanism
It is essentially 2 pieces of wire with weights on the ends. These wires are symmetrically wrapped around the equator of the rocket or the service/payload module (in our case) and the weights are secured by a release mechanism. After separation from the vehicle, at a preselected time, the weights are released, thus discarding enough momentum to reduce the spin of the module to desired values (generally of the order 0.1Hz). [4] Refer to Appendix 1.
Figure _ [5]: A stretch yo-yo consisting of weight, spring, wire and end fittings. (The scale is in inches)
Spin Stabilisation
Weighing in a number of spin stabilization systems such as a de-spun platform and 3-axis stabilization, the 'Yo-Yo' technique, after preliminary trade-off studies seemed to be the most feasible as it was shown to be the cheapest and lightest option. However, due to the nature of the technique the wires and masses will need to be replaced after each use which would contribute to an increased cost.
The chosen means for spin stabilization is a 'yoyo de-spinner'. The technique of yoyo de-spinning is by attaching 2 equal masses connected by wires of equal length to the exterior of the rocket by a pyrotechnic release mechanism. The wire is then wrapped around in the opposite direction to that of the rocket's spin. The principle behind this technique is that when the release mechanism has reached its predetermined time for release, the 2 masses are jettisoned tangentially from the rocket thereby increasing its moment of inertia and decreasing its angular velocity. Finally, once the distance of the masses from the rocket are equal to the length of the tether, they are then detached from the rocket and propelled off carrying a substantial fraction of the system's angular momentum.
For the proposed concept with an estimated mass of 300Kg, preliminary calculations put the masses to weigh a combined total of 3kg and a tether length of 10.05m when G=10-3.
Diameter
ω (initial)
ω (final)
I
L (tether)
0.3m
8.087Hz
0.2557Hz
6.75kg/m2
10.05m
0.8m
4.952Hz
0.387Hz
48kg/m2
10.05m
Using
- Where ω (final) = 0
Assuming that ω (final) equates to 0 despite early calculations should improve the final spin rate theoretically to 0. However as it will never be 0 exactly this improves the conditions for microgravity experiments by optimising the environment.
Sounding Rocket Avionics
Ground Control Computer/ Launch Station
Control the avionics system of the rocket and give direction and guidance command.
Packet Radio
Form of packet switching technology used to transmit digital data via radio/wireless communication links.
Flight Computer
Positioned in the rocket below nose cone. Acceleration, direction and the atmospheric pressure of the rocket is analysed. The data logged to the EEPROM (Electrically Erasable Programmable Read-Only Memory) for post launch and after launch analysis.
Control the parachute, rocket axis direction and GPS system.
GPS System
Interface with 3 Global Positioning System Satellite and constantly sent data to Flight Computer with is then analysed and send back to the launch station.
Accelerometer
Device used to determine the proper acceleration relative to the free fall by interface with the altimeter with used the ambient pressure and the pressure at that altitude to compute the temperature and pressure drop.
Sense the orientation, shock vibration, falling (Altimeter) and direction (axis). Axis models used to defect magnitude and direction as a vector quantity.
Constantly interface with the Flight Computer to interpret the data.
Altimeter
Pressure drop (barometer) and temperature drops logged in and interface with accelerometer.
Attitude Gyroscope
A gyroscope based inertially-referenced attitude sensor can be used to determine coarse 3-axis payload attitude information (1-3 degrees in all axes). An available option is the MIDAS which is manufactured by Space Vector Corporation. The MIDAS unit comprises of a pair of 2-degree of freedom displacement gyros, sensing the vehicle roll-yaw and pitch displacement respectively. [6]
INSTRUMENTATION POWER SUPPLY
The instruments on sounding rockets are driven by electrical power from batteries. Several types of battery system are available such as Silver Zinc cells and Nickel Cadmium cells which have a 3-4 times lower energy density than the former. There are two different types of the Silver Zinc cells, 'High Rate Discharge Series' designed for energy to be expended in an hour or less; and 'Manually Activated Primary Series' designed for quick activation.[7]
EVENT TIMING SYSTEM
Often time delays are desirable between various in-flight events, such as the motor separation and parachute deployment; or de-spin mechanism activation and research experiment activation. In-flight event timing is normally controlled by the following types of timers:-
Mechanical Timers:
These comprise of three basic components: "G" weight actuator, a spring wound timing mechanism, and an electrical switch system controlling external circuits, that are put into operation at the pre-set time. Three to eight switch units are available. Maximum time capacity ranges from 90 to 600 seconds.
Electromechanical Timers :
These are made up of a DC chronometrically governed timing motor and an electrical switch assembly controlling external circuits. Nine or thirteen switch units are available. And the maximum time capacity for electromechanical timers is 720 seconds.
Electronic Timers :
These are more complex timers with logic circuitry for low power consumption, a battery backup and time-event decoding programmes. Electronic timers can space events as close as 100 milliseconds apart. And can also provide a limited amount of random event programming. [8]
Figure 2.1 shows the schematic Avionics System block diagram of the sounding rocket which will be implemented at this initial stage of the project.
Figure 2.1
http://soundingrocket.org/Documents/ARES%20Rocket%20PPT-%202.pdf
Staging Process
1st Stage - launch station initiate launch sequence to flight computer and the ignition start off and open the valve. GPS and accelerometer activated on the valve is open.
2nd Stage - Pressure/Temperature difference is logged into the black box throughout the journey. Wireless link enable data transmission from the flight rocket to the launch station to prevent any loss of data.
3rd Stage - The separation mechanism is activated by accelerometer as the altitude and free falling condition is detected.
4th - The spinning motion of the payload module will initiate the 'YOYO' de-spin as the Angular-Rate-Gyroscope detects change in angular speed due to loss in weight.
Figure 2.2 shows the planned trajectory and staging of the sounding rocket.
Wireless Link Transmitter
5th Stage
1st Stage
#
2nd Stage
3rd Stage
4th Stage
Figure 2.2
Typical sounding rocket flight;
Credit: NASA
http://www.daviddarling.info/encyclopedia/S/sounding_rocket.html
Sounding Rocket Dimension
The importance of gaining sustainable method to determine the dimension is important as it plays major role of the entire project requirement and directly related to cost constraint. In order to obtain reliable source to determine the dimension of the sounding rocket, a systematic table been created by determining successful rocket of same kind. Table 1.1.1 shows the dimension parameter ratio of existing rocket meanwhile Table 1.1.2 shows the criteria range of the projects sounding rocket.
Dimension Parameter
Rocket Name
(A)
(B)
(C)
(D)
(E)
MT - 135
0.294
0.41
0.139
0.699
0.722
Maser
0.201
0.35
0.184
0.787
0.189
Black Brant (1-5)
0.221
0.43
0.243
0.978
0.226
MR - 12
0.222
0.31
0.231
0.919
0.474
TABLE 1.1.1
Parameter
(A)
(B)
(C)
(D)
(E)
(F)
Ratio
0.2-0.3
0.3-0.4
0.14-0.24
0.77-0.99
0.20-0.60
0.035-0.045
TABLE 1.1.2
(A) - Payload Height Fin Height
(B) - Diameter Height
(C) - Diameter Payload Height
(D) - Diameter Span
(E) - Fin Span Fin Height
(F) - Fin Height Apogee
With the help of this table, the initial design of the rocket was able to sketch and an allowable change in the dimension of 10% obtained. Even thought constrain such as technology and design implementation factors could change this ratio, at this stage of research this table can be applied to design the initial dimension. Figure 1.2 - 1.4 shows the initial sounding rocket dimension which been determined during the research.
Based on Table 1.1.1, the maximum allowable conceptual dimension of the sounding rocket is calculated and tabled on Table 1.1.3.
Dimension Parameters (cm)
Full Height
Diameter
Full Span
Payload Height
Fin Span
Fin Height
Table 1.1.3
RECOVERY SYSTEM
The main job of the recovery system is to bring the deliverable safely to the surface without damage. The most common Recovery system is the Parachute Recovery System.
Single Stage Recovery System
In this system a single parachute is used, which is put out at apogee. An altimeter, a timer or the motor sends an ejection charge for deployment.
Two Stage Recovery System
In this system a drogue parachute is put out at apogee and the main parachute is put out at a lower altitude like 600 feet. This is to minimize the drift in the rocket. An altimeter, a timer or the motor sends an ejection charge for deployment for both drogue followed by the main parachute.
Parachute Deployment Bag
It is a bag into which a parachute is put into. It might have more than one purpose, depending upon how the recovery system works. The two most common uses are to protect the parachute from hot ejection charge gases and particles, and for organised deployment. Nomex cloth is commonly used for making deployment bags because it is fire resistant.
http://www.info-central.org/images/263-002.jpg
Parachute Deployment Bag
http://www.info-central.org/images/263-002.jpg
Other Recovery Techniques
Gliding Recovery: Using Aerodynamic surfaces which can produce lift to control the terminal velocity. The aerodynamic requirements of gliding flight needs to have a mass shift to allow transition between vertical flight and gliding flight.
Featherweight recovery: If the rocket is very light (less than an ounce typically), or the rocket have enough drag that it's terminal velocity is very low.
Parachute Recovery: Using a parachute or parasheet for drag. Because of the efficiency of parachutes, this is the most popular way. You get more drag with less cloth than in any other way. Because of this efficiency they are used for virtually all high power projects.
Break-Apart recovery: In this way one could simply break the rocket in the middle and attach the 2 sections by a shock cord. Works for small rockets. It would be possible to make large rockets, with very large surface area and relatively low weight that would be safe to recover this way.
Helicopter Recovery: Rigid blades could be used to slow terminal velocity. For this system the whole rocket must be designed around this m the whole rocket is designed around it. The stresses of a rapidly spinning rocket touching down are enormous.
Streamer Recovery: The streamer adds drag and slows the rocket. Bigger streamer would be providing better recovery.
Hence considering all the alternatives, anything over a few ounces, parachute recovery is the most important, tried and tested, and very simple. They are the most efficient for their weight and bulk.
Aim of the Recovery System
The primary objective is to design a recovery system to recover the payload and deliverables along with the motor. The system will have to maintain structural and aerodynamic stability once it reaches the atmosphere to the surface, and survive impact.
Needs and Requirements:
Structural stability
30-60 sec free-fall or as long as possible
Parachute deployment system
It must endure natural flight conditions without rupturing.
It must spread reliably.
It must slow the deliverables enough to avoid ruinous damage to equipments.
The Deployment system must have different deployment triggers altitude triggers, or timed triggers and an emergency back-up trigger.
Constraints
Lightest weight possible
Smallest size possible
Communications
Low temperature since it might affect the internal instruments
Windy Conditions
Providing power for equipment
Money for best possible electronic instruments and structural materials
PARACHUTE DEPLOYMENT SYSTEM
The successful operation of any parachute depends on satisfactory performance of the deployment system. Deployment of a parachute is defined herein as the process of extending the parachute canopy and suspension lines to a position at which satisfactory inflation can occur.
Forced Ejection System is frequently used and a very common technique. This technique is very simple. The mortar, catapult, and pressure bellows are some of the techniques designed to provide a forced ejection of parachute.
Drogue Parachute System is parachute deployment using a drogue parachute and has numerous advantages. The system is quite flexible since the parachute extraction force is applied continuously over the entire deployment sequence. Compared with a forced ejection system, the drogue system is usually lighter in weight and produces smaller reaction loads. These reductions are due primarily to the reduced energy requirements of the drogue parachute mortar. A drogue system is more complex than a forced ejection system because the drogue parachute itself must be deployed.
Deployment Mechanisms
Following are comparative types of deployment system.
Spring Loaded
In this method a spring is used to deploy the parachute at the desired time.
Mortar
Typical ejection charge & flashbulb
In this method an electric signal is used to trigger the explosive charge in order the parachute could be blown out.
Drogue/Mortar
Typical closure ejection well
Once the drogue is released, it pulls the parachute out. The mortar is fired at the same moment to inflate the parachute following the drogue.
Parachute Deployment
Weight
a
b
C
d
Spring
Mortar
Drogue
Mortar+Drogue
Reliability
10
2
6
4
8
Size
6
2
3
4
1
Stability (Drift)
8
2
2
6
6
Trigger Integration
6
1
4
2
3
Cost
4
4
1
3
2
Simplicity
6
2
4
3
1
Total Score
82
146
154
166
Reliability is the most important aspect hence it gets a weightage of 10. Stability is the next important thing to look into. The cost is not so important since the material used to manufacture the parachute deployment system is comparatively light and less expensive. But the size, trigger integration and simplicity are fairly important hence getting a higher weightage.
PARACHUTE DESIGN
Following designs have been looked into:
Parachute with hole
Parachute with no hole
Parachute with a drogue and a hole
Parachute with a drogue but no hole
Parachute Design
Weight
A
b
C
d
with hole
no hole
with hole+drogue
no hole with drogue
Stability
10
3
1
4
2
Force of Deployment
5
2
1
4
3
Simplicity
10
3
4
1
2
Total Score
70
55
70
55
The design needs to be stable and simple hence getting a higher weightage, followed by the force of deployment.
Summary
The initial dimension are subjected to changes during the next stage of project but following the allowable constraint of 10% to avoid complex as well as drastic changes on the aerodynamic properties of the rocket. The constraint allowable limits also smoothes the designing process as the further trade of and other changes are limited to avoid more structural and cost spending.
While the avionics system are determined, the 'brain' of the sounding rocket now can be implemented into the rocket and adjustment to comply with the requirement and other essential systems needed to succeeding the launch can be completed.
Doing further research and using the outcome of the weighted decision matrix it could be safely concluded that a Dual Deployment system using a drogue and a mortar for deployment and a parachute with a hole is a feasible option.
some significant factors need to be measured: Sufficient clearance between the separating bodies; No damaging shock loads provoked in the structure and payload; No contamination such as extreme or destructive debris from the operation.
