Steven P. Powell

Signal Characteristics of Civil GPS Jammers

This paper surveys the signal properties of 18 commercially available GPS jammers based on experimental data. The paper is divided into two distinct tests. The first characterizes the jamming signals, and the second test determines the effective range of 4 of the jammers. The first test uses power spectra from discrete Fourier transforms (DFTs) of the time series data to show that all the jammers employ approximately Copyright ©2011 by Ryan H. Mitch, Ryan C. Dougherty, Mark L. Psiaki, Steven P. Powell, and Brady W. O’Hanlon, Jahshan A. Bhatti and Todd E. Humphreys. All rights reserved. Preprint from ION GNSS 2011 the same jamming method, i.e. linear frequency modulation of a single tone. The spectra also show that there are significant jammer-to-jammer variations, including between jammers of the same model, and that a given jammer’s signal may vary over time. The first test also includes measurements of signal power within frequency bands centered at the L1 and L2 frequencies, along with the sweep periods and the sweep range at both frequencies. The second test presents measurements of the attenuation of the jamming signal necessary to allow a commercially available GPS receiver to acquire and track signals from a GPS simulator. From the attenuation levels and some assumptions about the antennas used, upper limits on the effective jamming ranges are calculated for 4 of the jammers, with a resulting maximum range of 6–9 km.

Bit-wise parallel algorithms for efficient software correlation applied to a GPS software receiver

A set of efficient algorithms for processing code-division multiple-access spread spectrum signals has been developed. They make use of bit-wise parallelism to process 32 samples simultaneously. These algorithms have been implemented in a real-time global positioning system software receiver. The receiver consists of a radio-frequency front end, a system of shift registers, a digital data acquisition card, and software that runs on a 1.73-GHz PC. The PC performs base-band mixing and pseudorandom noise code correlations in a manner that directly simulates a hardware digital correlator.

Design and practical implementation of multifrequency RF front ends using direct RF sampling

The use of direct RF sampling has been explored as a means of designing multifrequency RF front ends. Such front ends will be useful to multifrequency RF applications such as global navigation satellite system receivers that use global positioning system (GPS) L1, L2, and L5 signals and Galileo signals. The design of a practical multifrequency direct RF sampling front end is dependent on having an analog-to-digital converter whose input bandwidth accommodates the highest carrier frequency and whose maximum sampling frequency is more than twice the cumulative bandwidth about the multiple carrier signals. The principle of direct RF sampling is used to alias all frequency bands of interest onto portions of the Nyquist bandwidth that do not overlap. This paper presents a new algorithm that finds the minimum sampling frequency that avoids overlap. This design approach requires a multifrequency bandpass filter for the frequency bands of interest. A prototype front end has been designed, built, and tested. It receives a GPS coarse/acquisition code at the L1 frequency and GPS antispoofing precision code at both L1 and L2. Dual-frequency signals with received carrier-to-noise ratios in excess of 52 dB-Hz have been acquired and tracked using this system.

First results from the Freja HF snapshot receiver

The Freja plasma wave instrument has measured electric field waveforms up to 4 MHz in the auroral ionosphere near 1700 km altitude. The HF snapshot receiver responds to natural signals during every passage through the auroral ionosphere and we have curren

Magnetometer-based Attitude and Rate Estimation for Spacecraft with Wire Booms

A magnetometer-based filter and smoother are presented for estimating attitude, rate, and boom orientations for a spinning spacecraft employing wire booms. The work is motivated by the need to estimate the attitude and boom configuration of a recent scientific sounding rocket mission. During flight, the primary payload ejected two small spinning spacecraft, or subpayloads. These each deployed four 3-m wire booms attached to small tip-mass probes. Following deployment, the booms were aligned approximately radially about the nominal spin axis. The attitude determination task is to reconstruct the attitude, rate, and boom orientations for each subpayload. The estimates are initialized by the approximate initial angular momentum for each subpayload at ejection, provided by inertial sensors onboard the primary payload, and thereafter are based on magnetic field data furnished by a three-axis magnetometer onboard each subpayload. The estimation process is complicated by the flexible wire booms whose full parameterization for even the simplest pendulous modes would require 16 state elements in addition to the 6 elements necessary to define the state of the main spacecraft body. In this work, several simplifying assumptions about the motion of the booms allow reduction of the problem to tractable complexity. A variant of an existing magnetometer-based attitude and rate estimator is developed which is suited to the time-varying model errors resulting from the simplifying assumptions. Two key features of this new estimator are its use of the inertial components of the angular momentum vector in place of the traditional angular rate vector in the state estimate and its explicit inclusion of an error model for the approximate relationship between angular rate and angular momentum. The estimator, a filter/smoother, is applied to synthetic data from a truth-model simulation and then to actual telemetry from the mission subpayloads. Although designed for a spinning spacecraft with flexible wire booms, the filter and its related smoother constitute good estimation algorithms for any spacecraft that must employ coarse Eulerian model-based estimation in the presence of significant modeling errors.

CASES: A Smart, Compact GPS Software Receiver for Space Weather Monitoring

A real-time software-defined GPS receiver for the L1 C/A and L2C codes has been developed as a low-cost space weather instrument for monitoring ionospheric scintillation and total electron content. The so-called CASES receiver implements several novel processing techniques not previously published that make it well suited for space weather monitoring: (A) a differencing technique for eliminating local clock effects, (B) an advanced triggering mechanism for determining the onset

Attitude Estimation for a Flexible Spacecraft in an Unstable Spin

Anattitudereconstruction algorithm hasbeendeveloped fora e exiblesounding rocketwhosespin vectornutates unstablyaboutitsminorinertia axis.Theattitudeestimatesareneededforposte ightanalysisoftherocket’ sscience data. An additional goal of the work is to show that a e exible-body model can be used in a Kalman e lter or in a smoother. The attitude is estimated using a smoother whose dynamic model includes a main rigid body and a pair of elasticbooms. Boom e exure is the principal causeof nutation instability. Boom bending is modeled by a Markov process, but the laws of mechanics are used to determine its ine uence on the attitude dynamics. This smoother achieves a better attitude estimation accuracy than can be achieved using a rigid-body model. The peak attitude error is estimated to be 4 deg, and the principal error source seems to be the limited accuracy of the rocket’ s attitude sensors. HIS work deals with the poste ight attitude determination for a sounding rocket mission, the Cleft Accelerated Plasma Experimental Rocket (CAPER). CAPER was launched from the Andoya Rocket Range in Norway in January 1999. Its e ight lasted slightly longer than 1200 s and reached an apogee altitude of 1360 km. The goal of this mission was to study the behavior of the ionosphere during auroral events. Attitude information is needed to transform CAPER’ s measurements of electric e eld components into geodetic coordinates. An attitude accuracy of from 2 to 4 deg is considered acceptable for the mission. From an attitude determination standpoint, the important characteristics of the CAPER sounding rocket were as follows: Its attitude sensors were a vector sun sensor with a slit e eld of view, a e xed-head horizon crossing indicator (HCI), and a three-axis magnetometer that was mounted on a short boom. The rocket’ s attitude was passively spin stabilized with the nominal spin vector directed along its minor inertia axis. CAPER deployed several booms after exit from the atmosphere and after the e nal stage motor burn. The longest of these were two e exible 3-m booms that deployed perpendicular to the nominal spin axis and in opposite directions from each other. A schematic diagram of this cone guration appears in Fig. 1. Minor-axis spin stabilization is often used in sounding rocket experiments. Near the dawn of the space age, the Explorer-1 mission demonstrated that such a cone guration has an unstable nutation mode due to energy dissipation in the e exible booms. 1 In many sounding rocket experiments it is acceptable for the nutation amplitude to grow as long as the resultant coning half-angle does not become too large by the end of the e ight. There are two major challenges in doing poste ight attitude determination for CAPER. The e rst is the lack of rate-gyro data. This challenge dictates the use of an Eulerian dynamics model to propagate the attitude and rates between measurement sample times. The model becomes especially important toward the end of the e ight, when only magnetometer data are available. The attitude estimate in rotation about the magnetic e eld vector is based solely on dynamic model propagation during this phase, and model inaccuracy

Rapid Energy Dissipation in a Yo-Yo-Type Wire Boom Deployment System

A wire boom deployment system has been developed that uses a mechanism that is similar to the classic yo-yo despin mechanism. The goal has been to develop a way to rapidly deploy wire booms from a spinning soundingrocket experiment. The main challenge in using a yo-yo-type mechanism is to dissipate the excess kinetic energy so that the wire booms do not rewrap themselves about the spacecraft after deploying. A ring has been added to the basic yo-yo mechanism. It rotates with respect to the main spacecraft about the nominal spin axis, and the wire boomsdeploy from it. Thering/spacecraftjoint isdamped, and relativemotion betweenthering and thespacecraft dissipates the excess energy. This system is analyzed and simulated in two and three dimensions. The results show that a well-tuned system can deploy 2.5 m wire booms in under 20 s and that it can tolerate the expected levels of asymmetry and parametric uncertainty.

Practical Design and Flight Test of a Yo-Yo Wire-Boom Deployment System

Ay o-yo-type wire-boom deployment system has been developed and flight tested on a sounding rocket mission. The goal of the work has been to validate a new mechanism that rapidly deploys wire booms from a spinning spacecraft. This work takes a theoretical system design and implements it in practical hardware. The limitations inherent in practical hardware necessitated new theoretical developments. A modified stability analysis has been developed for the case of nonzero axial separation between the wire-boom base attachment points and the spacecraft’s center of mass. This modified stability analysis dictates that a stable design is impractical for many missions because very large wire-boom tip masses are needed and because the three-dimensional deployment transients are very sensitive to small asymmetries. This problem has led to the development of design criteria and analysis techniques, which permit a short-duration mission to use a slightly unstable nutation mode. These techniques have been used to design a system that has been flown on two daughter spacecraft that were part of a formation of three sounding rocket subpayloads. Each daughter spacecraft deployed four 3-m-long wire booms in under 10 s and maintained a low level of spin instability for the remaining 700 s of the mission. The nutation oscillations showed a slow exponential growth, but the coning half-angles of both spacecraft never exceeded 16 deg.

Searching for Galileo

BIOGRAPHIES Mark L. Psiaki is a Professor in the Sibley School of Mechanical and Aerospace Engineering. He received a B.A. in Physics and M.A. and Ph.D. degrees in Mechanical and Aerospace Engineering from Princeton University. His research interests are in the areas of estimation and filtering, spacecraft attitude and orbit determination, and GNSS technology and applications. Todd E. Humphreys is a graduate student in the Sibley School of Mechanical and Aerospace Engineering. He received his B.S. and M.S. degrees in Electrical and Computer Engineering from Utah State University. His research interests are in estimation and filtering, spacecraft attitude determination, GNSS technology, and GNSS-based study of the ionosphere and neutral atmosphere.