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The electronics of the H6 rocket



Instrumentation

The instrumentation of the H6 rocket consists of several units which collect measurement data:

  • a pressure sensor,
  • a temperature sensor on the tip of the nose cone
  • a vibration sensor
  • a thrust sensor
  • an accelerometer with analogue integrator
  • a magnetometer (provided that there is enough time to implement this)

Recovery timers

The recovery system of the H6 rocket is completely autonomous and contains everything that is needed to guarantee a safe recovery during the flight. This includes two timers and the corresponding pyrotechnical devices (for the parachute hatch and for the release of the main parachute). The timers are digitally adjustable. The basis for the design of the timers and the pyro ignition circuitry is simplicity and reliability. An effort has been made to rule out the failure we had experienced with the previous rocket, the H5. The timers are to a large extent insensitive to malfunctions. The recovery system and the instrumentation of the H6 rocket are almost completely separated from each other. Only the battery power supply is shared between the systems, although even this could be separated, but this was not considered to be necessary.

Status information

Apart from the instruments for external measurements, several rocket system parameters are also recorded, such as internal temperatures, stress levels, uplink signal strength and 24 (maximum) status parameters. The analogue signals are selected and sampled by an analogue multiplexer and subsequently fed into an 8 bit analogue/digital  (A/D) converter. After conversion, the data are mainly stored in non-volatile RAM. Only a fraction of the data are transmitted in real time to the ground station. The most important for this is that the downlink has a limited capacity, just over 50 Bytes/s. Within the H6 rocket about three to four times this amount of data is produced.

With the H5 rocket experiences at the back of our minds, special attention has been paid to the memory for data storage. For this purpose, a so-called Impact Resistant Module (IRM) has been designed. As the name implies, this module should be able to survive an impact of the rocket, without resulting to damaged to the equipment inside.

Microcontroller

The micro controller, which forms the hart of the data acquisition system is of the type 68HC11. Practically all the functions required for this purpose are integrated on the chip itself. It contains 12 K EPROM (behind a frame), 512 Byte EEPROM and even 512 Byte RAM, which, provided that an external back-up battery has been mounted, will not be erased in the case of loss of power of the main battery. The 68HC11 is equipped with two serial bus interfaces:

  1. RS-232, which is used for the up/down link,
  2. Serial-Parallel Interface (SPI) bus, for the data input/output.

Furthermore, an 8 channel (8 bit) A/D converter is present on the chip. This one is not used because all A/D conversions are carried out on the analogue modules themselves, while the data subsequently flows via the SPI bus to the controller. There is just one exception to this data flow, viz. the status data of the 5 V power supply for the digital modules.

All relevant flight data is stored in RAM. such that short lasting power losses do not disrupt a proper function of the system. In case of a power loss longer than 3 seconds. a system reset is generated and the controller is put into the pre-lift-off condition. The lift-off signal itself is received from the timer module on which the g-switch and the safe/arm circuit are mounted.

The power consumption of the controller can be limited since the system is able to go into a WAIT state. The 68HC11 has a completely static design, which means that every clock speed between 0 and maximum (about 2 MHz) is acceptable. It is expected that the complete Printed Circuit Board (PCB), including the interface and several special circuits have a power consumption of about 30 to 40 mA.

Accelerometer

An important part of the instrumentation is dedicated to do measurements which will lead to determining the relation between the Mach number and the air drag coefficient. For this it is necessary to know the acceleration and the velocity of the rocket at every moment (and especially during the first 30 seconds ) of the flight. The acceleration is measured with a solid-state accelerometer of the type IC-sensors, model 3021 (50 g range). By integrating the acceleration analogously on-board the rocket, the actual velocity of the rocket is calculated.

The analogue integrator is started by the LIFT-OFF signal coming from the timer module. Both the acceleration and the velocity are sampled at a frequency which depends on the flight phase. Before lift-off, this can be as low as 2 Hz. Directly after lift-off, during the powered flight, the sampling rate has to be significantly higher in order to measure the behaviour of the motor. During this phase of the flight, violent vibrations in the structure can disturb the measurement of the acceleration. The output signal of the analogue integrator shows much smaller fluctuations and the maximum value of this signal (just after cut-off of the motor) gives a good indication of the velocity of the rocket.

In order to determine the maximum altitude of the rocket flight, the acceleration of the rocket could be integrated twice, yielding the displacement. However, the altitude can be determined more accurately using the on-board pressure and temperature sensors (P/T sensors). By numerically integrating the measured pressure vs. temperature profile of the atmosphere, the altitude can be determined as a function of the pressure. The NERO rockets Interim-03 and Pollux have flown in 1986 and 1987 with on-board P/T sensors and the results from these flights were encouraging. The actual P/T profile measurement is performed during the descent of the rocket by parachute. During the ascent of the rocket, when the velocities of the rocket are much higher, interesting (temperature) measurements can be carried out on the top of the nose cone of the rocket. Because of adiabatic compression, the temperature at this point can increase several tens of degrees Celsius.

Vibration sensors

The final experiment on-board the H6 rocket is the measurement of vibrations caused by the motor. After analysing the H5A flight data, we suspected that the level of the vibrations inside the rocket were very high. The goal of this experiment was to be able to estimate the order of magnitude of the vibration levels inside the rocket in the future. For this purpose, the displacements of certain representative structural parts of the rocket were measured by using strain gauges. However, the amount of data in the vibration signal proved to be so large and the measurement duration was so small (several seconds) that is was not possible to determine a complete time or frequency domain graph of the vibrations. Therefore, only the energy (or average amplitude) in a certain frequency band is measured. For this, three frequency bands were selected on theoretical grounds which would contain the major part of the vibrational energy. The problem however, is how to interpret the measurement data, since the vibration level is not only dependent on the vibration source (i.e. the motor), but also on the properties of the structure itself. Furthermore, acoustic coupling between the structural parts of the rocket might also play an important part. At this moment, the final word in this matter has not yet been spoken. It is most likely that the vibration sensor will be mounted on a PCB.


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101KB
Take a look at the lay-out of the electronics of the H6 rocket.


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