Russia Adv – 1ac

***Solvency Stuff

Solvency – 1ac

SRGs enable us to get to Mars and beyond with limitless energy – DAs are non-unique because RTG’s in the status quo are worse

NASA 2K2 – (Space Radioisotope Power Systems Stirling Radioisotope Generator official NASA REPORT, pg 1-2, http://www.aboutnuclear.org/docs/space/stirling.pdf)
Radioisotope power systems can provide continuous power for 20-plus years, and have been used safely and reliably over the past 30 years in regions of space where the use of solar power is not feasible. To date, the United States has launched 25 missions involving 44 Radioisotope Thermoelectric Generators (RTGs). Although Stirling Radioisotope Generators have not been launched in a space exploration mission to date, they too are an application of a radioisotope technology that is well understood. To enable the next ambitious steps in exploration of our Solar System with safe, cost effective spacecraft, the U.S. Department of Energy (DOE) and National Aeronautics and Space Administration (NASA) are developing advanced, high-efficiency radioisotope power converters. The Stirling Radioisotope Generator (SRG) is one of the technologies being developed to provide spacecraft onboard electric power for potential use on future NASA missions. The development of the SRG will build upon a 55-watt-electric Stirling convertor previously developed under DOE contract with the Stirling Technology Company (STC), with NASA Glenn Research Center (GRC) assistance. The efficiency of the Stirling convertor was demonstrated to be in the mid Space Radioisotope Power Systems Stirling Radioisotope Generator April 2002 20 percent range. Use of the 55-watt-electric convertor in an SRG could reduce the required amount of radioisotope fuel (plutonium-238 dioxide), thereby potentially reducing the cost and amount of plutonium-238 dioxide flown on future missions. How Does a Stirling Convertor Work? The 55-watt Stirling convertor is a free-piston machine that operates on a Stirling thermodynamic cycle. Heat is supplied to the convertor from a DOE General Purpose Heat Source (GPHS) module, containing approximately 600 grams of Plutonium dioxide, and producing about 250 watts of thermal power. The heat input to a convertor results in a hot-end operating temperature of 650oC. Heat is rejected from the cold end of the convertor at nominally 80oC. The closed-cycle system converts the heat from a GPHS module into reciprocating motion with a linear alternator resulting in a AC electrical power output of 60-62 watts. An AC/DC convertor in the Stirling convertor controller converts the AC power to approximately 55 watts DC. What Are the Current Development Plans? The need for safe, reliable, long-lived power systems for future missions includes surface exploration of planetary bodies such as Mars as well as missions in the vacuum of space beyond Earth orbit. DOE and NASA are initiating the development of a Stirling Radioisotope Generator (SRG) power system that could be used for a variety of missions. The design goals for the SRG include ensuring a high degree of safety, optimizing power levels over a minimum lifetime of 14 years, and minimizing weight. The SRG will be designed to operate on planetary bodies as well as in the vacuum of space. In addition, it will be designed to deliver 100 to 120 watts of DC electric power. Each SRG will utilize two 55-watt Stirling convertors with about 500 watts of thermal power supplied by using two GPHS modules.

Solves Mars Colonization

SRGs get to mars

Thieme and Schreibe 2005

[Lanny G and Jeffrey G. National Aeronautics and Space Administration Glenn Research Center, “Supporting Development for the Stirling Radioisotope Generator and Advanced Stirling Technology Development at NASA Glenn Institute” http://www.bitecsmi.com/NASAdevelopment-213409-2005.pdf]
A high-efficiency, 110-We (watts electric) Stirling Radioisotope Generator (SRG110) for possible use on future NASA Space Science missions is being developed by the Department of Energy, Lockheed Martin, Stirling Technology Company (STC), and NASA Glenn Research Center (GRC). Potential mission use includes providing spacecraft onboard electric power for deep space missions and power for unmanned Mars rovers. GRC is conducting an in-house supporting technology project to assist in SRG110 development. One-, three-, and six-month heater head structural benchmark tests have been completed in support of a heater head life assessment. Testing is underway to evaluate the key epoxy bond of the permanent magnets to the linear alternator stator lamination stack. GRC has completed over 10,000 hours of extended duration testing of the Stirling convertors for the SRG110, and a three-year test of two Stirling convertors in a thermal vacuum environment will be starting shortly. GRC is also developing advanced technology for Stirling convertors, aimed at substantially improving the specific power and efficiency of the convertor and the overall generator. Sunpower, Inc. has begun the development of a lightweight Stirling convertor, under a NASA Research Announcement (NRA) award, that has the potential to double the system specific power to about 8 We /kg. GRC has performed random vibration testing of a lower-power version of this convertor to evaluate robustness for surviving launch vibrations. STC has also completed the initial design of a lightweight convertor. Status of the development of a multidimensional computational fluid dynamics code and high-temperature materials work on advanced superalloys, refractory metal alloys, and ceramics are also discussed. Under the auspices of NASA’s Prometheus project, the Department of Energy (DOE), Lockheed Martin (LM) of Valley Forge, Pennsylvania, Stirling Technology Company (STC) of Kennewick, Washington, and NASA Glenn Research Center (GRC) are developing a high-efficiency, nominal 110-We (watts electric) Stirling Radioisotope Generator (SRG110) for possible use on future NASA Space Science missions. The SRG110 is being developed for multimission use (e.g., in operating environments with and without atmospheres); potential missions include providing electric power for unmanned Mars rovers and deep space missions. The SRG110 would provide a highefficiency power source alternative to Radioisotope Thermoelectric Generators (RTGs). The SRG110 system efficiency of greater than 20 percent would reduce the required amount of radioisotope by a factor of four or more compared to RTGs. LM, under contract to DOE, is the System Integration Contractor for the SRG110. LM is now developing the SRG110 Engineering Unit and will soon be proceeding on the Qualification Unit. The SRG110, described by Cockfield and Chan (2002), is expected to produce at least 112 We at beginning-of-mission (BOM), using two opposed Stirling convertors and two General Purpose Heat Source (GPHS) modules. The system efficiency is projected to be 22 to 25 percent with a system mass of less than 34 kg. STC is developing the Stirling convertor for the SRG110. This convertor was originally known as the Technology Demonstration Convertor (TDC). A total of 16 TDCs have been built by STC. The latest four were built with additional quality assurance practices that STC has implemented to prepare for flight convertor fabrication.
NTR is key to space propulsion

Kharytonov et al 11 (Oleksii M*, Boris M. Kiforenko**, Taras Shevchenko National University of Kyiv, 64, Volodymyrska St, 01033 Kyiv, Ukraine* S.P. Timoshenko Institute of Mechanics NASU, 3, Nesterov St, 03057 Kyiv, Ukraine**

(4-11, “Finite-thrust optimization of interplanetary transfers of space vehicle with bimodal nuclear thermal propulsion”ScienceDirect)
Three classes of the nuclear propulsion systems are considered as the promising technologies for primary space propulsion, namely the nuclear thermal rocket (NTR) engines (high-thrust propulsion), nuclear electric propulsion (NEP) (low-thrust propulsion), and bimodal nuclear thermal rocket (BNTR) engines, which can be used in both high and low thrust modes. The advantages of the NTR (high engine thrust levels (10–1000 kN) in comparison with 10–1000 N for the NEP) and the NEP (high specific impulse (more than 2000 s) in comparison with 1000 s for the NTR) are joined in the BNTR propulsion systems. So, the BNTR makes it possible to carry out the manned missions with both the small transfer duration and the high payload mass [1]. For the transfers of the vehicles with BNTR the problem of optimization of both the combination of highand low-thrust arcs and the parameters of NTR and NEP subsystems must be considered. The interplanetary trajectory of the SC is formed by the high-thrust NTR burns, which, as a rule, define a planet-centric maneuvers and by the low-thrust heliocentric arcs, where the NEP is used. Lower Dv 2 budget (Dv is velocity change) for the highthrust maneuvers corresponds to the higher final payload mass. The high-thrust arcs have to be analyzed by using a finite-thrust approach instead of a traditional impulse Nomenclature A coefficient in NEP specific mass model a jet acceleration a0 initial thrust-to-weight ratio of spacecraft B coefficient in NEP specific mass model b coefficient in the expression for new independent time variable C coefficient in NEP specific mass model c dimensionless constant in expression for gravity acceleration Cp average specific heat D dimensionless constant in expression for reactor thermal power E eccentric anomaly e eccentricity F coefficient in NEP specific mass model f the absolute value of Laplace vector G functional of Mayer-type variational problem g gravity acceleration gn scale of gravity acceleration H Hamiltonian h coefficient in the expression for new independent time variable i index number J integral functional j index number k index number L dimensionless constant in boundary conditions m mass ml (t) current value of SC mass at heliocentric lowthrust arc mp payload mw propulsion system mass mn scale of mass N reactor thermal power Nel electric power Nn scale of power n coefficient in NEP specific mass model P engine thrust Pn scale of engine thrust p parameter of Keplerian orbit q propellant mass flow rate qn scale of propellant mass flow rate R radius of circular heliocentric orbit of planet r radius-vector r0 radius of circular planet-centric orbit rp radius of pericentre of hyperbolic orbit rn scale of linear dimensions T propellant temperature Tn scale of temperature t time tn scale of time V exhaust velocity ! v vector of spacecraft velocity ! v 1 vector of spacecraft velocity on the boundary of sphere of influence ! v orb vector of heliocentric velocity of the planet vn scale of velocity x coordinate y coordinate a low-thrust NEP subsystem specific mass kg/kWt b tank coefficient g coefficient in the expression for exhaust velocity D change in parameter Z coordinate in transporting reference frame y the angle of thrust direction W true anomaly k dimensionless parameter in the expression for the SC final mass in the heliocentric arc l high-thrust NTR subsystem specific mass, kg/kN m gravity constant n the constant of proportionality between the maximal and minimal values of reactor thermal power x coordinate in transporting reference frame t new independent time variable c adjoint variable o0 angular position of SC in the circular planetcentric orbit 2 See Nomenclature. 224 O.M. Kharytonov, B.M. Kiforenko / Acta Astronautica 69 (2011) 223–233(burn) approach with allowance for the gravity losses. That is because of the burn duration is long (up to 45 min for trans-Mars injection [1]). It should be noted that the finite-thrust optimization of the planet-centric maneuvers is traditionally accomplished independently [2]. However, for the bimodal interplanetary transfer optimal distribution of the propellant budget between high- and low-thrust arcs depends on the maneuver conditions and the propulsion system parameters. So, one combined optimal control problem must be solved to optimize the motion at all high- and low-thrust arcs. The corresponding dynamical system has discontinuous right parts and its phase space changes in the instants of high- and low-thrust arcs conjunction.