Fiber optic gyro space application technology

Fiber optic gyro technology for space

1 High Reliability Technology

A typical digital closed-loop interferometric fiber optic gyro consists of a light source, coupler, Integrated Optic Chip (IOC), fiber optic ring, photodetector and signal processing circuit,The IOC and fiber optic ring form the sensitive loop of Sagnac interferometer, which is used for sensitive tacho information.The IOC and the fiber ring form the sensitive loop of Sagnac interferometer, which is used for sensitive rotational speed information.When the rotational speed in inertial space has an axial component Ω in the sensitive loop of the interferometer, a phase difference is generated between two beams propagating in opposite directions in the sensitive loop ΔØs ^[1]. According to the Sagnac effect, the relationship between the phase difference and the input rotational speed can be expressed as:

where D and L are the diameter of the sensitive loop and the length of the fiber, respectively, λ is the optical wavelength, and C is the speed of light. The two beams propagating in opposite directions in the sensitive loop The two beams of light propagating in opposite directions in the IOC interfere with the output, and the detector converts the interfering light signal into an electrical signal, which is finally demodulated in the signal processing circuit to output the speed value.The detector converts the interferometric light signal into an electrical signal and finally demodulates the output speed value in the signal processing circuit. By applying a negative feedback signal to the IOC,Realization of closed-loop control of interference phase operating points ^[9].

Fig. 1 Diagram of single axis IFOG.

However, the light source, photodetector, fiber optic ring and signal processing circuit are susceptible to radiation, temperature and other space environments.The failure modes and failure rates of each component of the fiber optic gyro in the space environment are listed in Table 1 ^{[10] [11]}.

Fig. 2 Configuration of six-independent-axis IFOG

To improve the reliability of the fiber optic gyro, a six-axis redundant configuration is a more traditional design solution, as shown in Figure 2. This scheme achieves a backup of all optoelectronic devices, eliminating a single point of failure of the component. However, such a simple redundancy lacks However, such a simple redundancy lacks specificity and increases reliability at the expense of exponential growth in mass and power consumption, limiting the application of fiber optic gyros on satellites. To solve this problem, a four-axis fiber optic gyro configuration with shared dual light sources can save two gyros while still achieving The configuration is shown in Figure 3. As shown in Figure 3, this configuration consists of two light source modules and a four-axis fiber optic gyro.The two light source modules contain independent light sources and drive circuits, which are connected to the four-axis fiber optic gyro through two 1×4 couplers, each axis gyro contains independent optical and signal processing circuits, and the whole set of control circuits is used to realize interface communication, status monitoring, fault diagnosis and recovery control. The X, Y and Z-axis gyros are set orthogonally to each other for sensitive angular velocity in X, Y and Z directions, and the redundant S-axis is set at a certain angle tilt, so that when any one of the other three gyros fails, the S-axis will be activated as a replacement. When the primary light source module fails, the second light source module will be enabled as a replacement. Compared with the traditional stand-alone six-axis redundant configuration, this solution reduces the two-axis backup gyro with four light source modules, resulting in a significant reduction in component weight, power consumption and envelope size.

Fig. 3 Configuration of four-axes with two sharing sources IFOG.

The switchover of the backup component depends on the real-time status monitoring and fault diagnosis of the fiber optic gyro, based on the optical power and temperature parameters purely software-based fully digital in-orbit real-time monitoring method ^[12][13], the in-orbit fault diagnosis of the fiber optic gyro can be achieved by monitoring the in-orbit random wander The on-orbit fault diagnosis of fiber optic gyro can be achieved by monitoring the on-orbit random travel coefficient. First, for the space-use fiber optic gyro, the random travel coefficient can be estimated by the following equation.

Where k is the Boltzmann constant, T  is the temperature, e  is the electronic charge, Dv is the spectral bandwidth of the light source, R  is the denotes the detector transimpedance, I_d denotes the detector dark current, and P  denotes the optical power value of the system. Based on the estimation method of random wandering coefficient and the effect of space radiation and other factors, the prediction model of random wander coefficient of fiber optic gyro for space use can be expressed as ^[14][15].

Where, h  denotes the responsiveness of the detector, P₀ denotes the power of the light source coupled into the optical path,  A_c denotes the total optical path loss generated by the fiber coupler, the integrated optical path and the fusion part; q、b、f  is the fiber radiation sensitivity constant, d is the radiation dose,r is the radiation dose rate. Max(RWC_p) is the maximum value of RWC_p in the whole space mission under normal conditions, and the normal operation of the gyro satisfies.

RWC_e=Max(RWC_p)                        (4)

If RWC_e>Max(RWC_p),Then the gyroscopic optical path is determined to be faulty. The random wandering coefficient of in-orbit data is calculated in real time using the Allan variance RWC_a ，When the deviation of the on-track calculated value of the random wandering coefficient from the estimated value of the random wandering coefficient is greater than the set threshold value when the ，That is RWC_a – RWC_e =Threshold , then it can be determined that the gyro circuit is faulty. The diagnosis flow is shown in Figure 4.

Fig. 4 Flowchart of IFOG fault diagnosis.

2 Miniaturization technology

For the background of high reliability and long life application requirements, the dual light source four-axis integrated fiber optic gyro design scheme is a typical and efficient configuration scheme. efficient configuration, but the overall redundancy of the sensitive axis components still makes the weight and power consumption of the fiber optic gyro assembly based on this solution high. Power consumption is high. In recent years, the low-orbiting micro-nano-satellite application scenario has become increasingly popular, which requires further lightweight and low power consumption of fiber optic gyro. The recent low-orbit micro-nano-satellite application scenarios have raised higher requirements for further lightweight and low power consumption of fiber optic gyro.

In response to the above problem, a light and small three-axis integrated fiber optic gyro configuration scheme is shown in Fig. 5(a), where the orthogonally set The three-axis gyro is connected to the light source module through a coupler. In addition, the use of time-division multiplexing technology can further realize the miniaturization of the fiber optic gyro. The time-division multiplexing-based design in Figure 5(b) is the simplest configuration of the current three-axis integrated fiber optic gyro. Based on Fig. 5(a), this scheme again simplifies the composition of the fiber optic gyro, using only one coupler and one photodetector, reducing the complexity of the optical path and the number of fiber fusion points, and using only one set of signal processing circuit, which greatly reduces the quality and power consumption of the gyro assembly. The signal processing circuit adopts time-division multiplexing technology, during the gyro operation, the three axes work alternately in a certain order, By applying appropriate modulation to the three axial IOCs so that only one axial gyro works in the active state at each moment, the digital signals at each moment are demodulated, integrated and fed back in the FPGA, thus realizing the digital closed loop of the three-axis time-division multiplexed fiber optic gyro. Since the failure rate of the light source is the highest in space applications, the above two configuration schemes can improve the reliability by adding backup light sources.

Fig. 5 Configuration of (a)miniature 3-axis IFOG and (b) 3-axis IFOG based on TDM technology.

The time-division multiplexing technique is able to achieve the minimal structure of multi-axis integrated fiber optic gyro by further multiplexing of optical paths and circuits, but it also brings many problems, such as loss of information, deterioration of dynamic performance due to serial processing of data in each axis, and inter-axis crosstalk ^[16][17][18]. While the hardware structure is simplified, the complexity of the software is greatly increased, which also increases the development difficulty of the time-division multiplexed fiber optic gyro. For the N axis integrated time-division multiplexing fiber optic gyro, the sample capacity of the gyro output data is reduced to 1/N, the original one, during the period when all N axes are closed-loop working once.This degrades the limiting accuracy to a non-divided state ,And the transition process during the switching of each axis makes the amount of information shrink again, causing the time-division multiplexed fiber optic gyro in dynamic characteristics of the disadvantage ^[19]. However, considering the size, power consumption, cost and accuracy, and dynamic characteristics of fiber optic gyro compromise the time-division parameters, the time-division multiplexed fiber optic gyro still has a strong competitive advantage in low to medium accuracy space applications. The time division multiplexed fiber optic gyroscope still has a strong competitive advantage in low to medium accuracy space applications. The optical signal detected by the time-division multiplexed fiber optic gyro detector comes from the superposition of the signals of each sensitive loop, which makes the detector The scattering noise of the detector and the intensity noise of the light source become the main factors affecting the accuracy of the gyro, and the selection of a suitable modulation and demodulation scheme can A suitable modulation and demodulation scheme can reduce the degradation of signal-to-noise ratio caused by crosstalk noise ^[20][21][22].

In addition to this, miniaturization of optoelectronic devices is also an important means of gyro miniaturization. With the same configuration scheme, fine diameter fiber, small light source, small IOC, integrated PIN-FET, optimized small integrated digital processing circuit, etc. can significantly The size and mass of the fiber optic gyro can be significantly reduced. Figure 6 shows some of the fiber optic gyro miniaturization devices, using a fine diameter 100um coated layer The average diameter of the wound fiber ring can be reduced to 19 mm using a fine-diameter, 100-um deflection-preserving fiber.

Fig. 6 Miniaturized components: (a)PIN-FET, (b)Small-diameter PMF, and (c)Mini fiber coils.

3 Irradiation resistant technology

Satellites are exposed to harsh space radiation during orbital operation, which causes radiation-induced attenuation (RIA) of the fiber optic.
Radiation-induced attenuation (RIA) seriously affects and limits the accuracy of gyroscope, especially for satellites in high orbit, the effect of irradiation is more prominent, so it is necessary to further improve the performance of fiber optic gyroscope against irradiation. The advantages of photonic crystal fiber (PCF), which was proposed by Philip Russell in 1991 and made for the first time at the University of Southampton in 1996, have been proven from theory and engineering, and the radiation-induced attenuation of polarization-maintaining photonic crystal fiber (PM-PCF) is tens of times lower than that of commonly used polarization-maintaining fibers. The design of photonic crystal fiber can substantially improve the irradiation resistance of gyroscopes [23]. However, its application in gyroscopes has been limited due to high transmission loss, high fusion difficulty, and poor fusion point reliability.

Fig. 7 (a)Cross Section of the PM-PCF and (b) the PM-PCF coil.

After years of research, in 2017, the low-loss fine diameter PM-PCF developed by Beijing University of Aeronautics and Astronautics broke through the transmission loss to 1.1 dB/km, which can meet the application requirements of high-precision fiber optic gyroscope, the cross-section of the fiber is shown in Figure 7(a), and the structure The cross-section of the fiber is shown in Figure 7(a), and the structure and performance parameters are shown in Table 2. Figure 8 shows the RIA of PM-PCF at different optical wavelengths under the condition of 500Gy total radiation dose. The RIA of the PM-PCF at different wavelengths under 500Gy total radiation dose is shown in Fig. 8. The performance is improved by an order of magnitude compared with the conventional bias-preserving fiber ^[24].

The fiber ring wound with the above photonic crystal fiber is shown in Fig. 7(b), and the ring diameter is 120 mm. The zero-bias stability of the developed PM-PCF fiber gyro is better than 0.002°/h, and the specific parameters are shown in Table 3. The specific parameters are shown in Table 3.

Typical fiber optic gyro products for space

Table 3 shows the typical space-use fiber optic gyro products and their performance parameters developed for different space mission requirements, and the ground test results of zero offset and noise are shown in Figure 9. The “miniaturized four-axis dual-light fiber optic gyro” and “miniaturized three-axis fiber optic gyro” have been working smoothly on many satellites in orbit as micro-compact, highly reliable and mature fiber optic gyro products. Polarization-preserving photonic crystal fiber optic gyro has been applied to cargo spacecraft, and the use of polarization-preserving photonic crystal fiber optic solution is an important trend in the development of fiber optic gyro for space applications.

Table 3 IFOGs for satellites developed by Beihang University

Model	Miniature Four-axis IFOG	Miniature Three-axis IFOG	PM-PCF IFOG
Appearance
Bias stability （100s, 1σ）	≤0.02°/h	0.2°/h	≤0.002°/h
Random walk coefficient	≤0.005°/√h	0.05°/√h	≤0.0004°/√h
Scale factor performance	≤ 50ppm	≤100ppm	≤ 10ppm
Dimension	135mm×135mm ×100mm	80mm×50mm ×50mm	—
Mass	≤2kg	≤0.4kg	—
Power consumption	≤10W	≤4W	—