Accounts Sba Essay Example for Free

Accounts Sba Essay The aim of my project is for me to arrive at a comprehensive understanding of the financial sector of the business world. To draw up the financial records for the year 2009 for Jen’s Supermarket and to comment on the entity’s performance using appropriate tools of Analysis Description and Overview Jen’s Supermarket is a developed and simulated entity owned by Jenevonne Kirton, who is the sole proprietor. Jen’s Supermarket is a developed supermarket which sells a variety of food items and miscellaneous goods, serving most of the island in its convenient location. The supermarket employ fifteen staff along with Ms. The capital at beginning was 202800 while at end it was 208165 which showed an increase by 5365 at the end of the month. Suggestions/Recommendations Current Ratio It is recommended that Jen’s Supermarket 1. Uses the excess money to expand the business by opening another supermarket. 2. Introduce new product to the business and purchase new equipment. Gross Profit It is suggested that Jen’s Supermarket should 1. Source cheaper suppliers to facilitate the demand of goods so profitability would be greater at the end of the month. 2. Offer new promotions every week and offer the customers raffles at a chance to win prizes at the end of the year. These are given to help Jen’s Supermarket increase gross profit and sales. Conclusion It is clear that Jen’s Supermarket keeps adequate records for the business to demonstrate the profitability over the financial period for the month. It could however, increase the current ratio and gross profit by incorporating the suggestions/recommendations outlined. Nevertheless Jen’s Supermarket made a good net profit for the month of December in 2009. Overall based on the performance of the business Jen’s Supermarket was able to make a profit of 6365 despite the low sales of 104 850.

Foreign exchange Essay Example for Free

Foreign exchange Essay As mentioned above, there was a significantly high supply of money in the United States, which characteristically led to depreciation of the United States dollar. In addition, the period was characterized by a high rate of inflation, implying that considerably larger sums of money would be used in buying commodities that would have been bought inexpensively in the earlier period (Atkinson and Hutto). On the international scale, the high cost of United States commodities meant that importers had to pay lots of money to acquire the goods, which ultimately lowered the level of international trade between the United States and other countries (Table 3). Consequently, the United States dollar fluctuated against other world currencies such as the Japanese yen, the Sterling Pound, Chinese yuan and the Australian dollar (Atkinson and Hutto). For instance, the Sterling Pound exchanged at between $1. 9548 and $2. 0442 in 2007 (Table 3), which was a very large variation. period Australia (dollar) Canada (dollar) China, P. R.(yuan) EMU Members (euro) Japan (yen) Mexico (peso) South Korea (won) Sweden (krona) Switzerland (franc) United Kingdom (pound) 2000. .5815 1. 4855 8. 2784 . 9232 107. 80 9. 459 1,130. 90 9. 1735 1. 6904 1. 5156 2001 . 5169 1. 5487 8. 2770 . 8952 121. 57 9. 337 1,292. 02 10. 3425 1. 6891 1. 4396 2002 . 5437 1. 5704 8. 2771 . 9454 125. 22 9. 663 1,250. 31 9. 7233 1. 5567 1. 5025 2003 . 6524 1. 4008 8. 2772 1. 1321 115. 94 10. 793 1,192. 08 8. 0787 1. 3450 1. 6347 2004 . 7365 1. 3017 8. 2768 1. 2438 108. 15 11. 290 1,145. 24 7. 3480 1. 2428 1. 8330 2005 . 7627 1. 2115 8. 1936 1. 2449 110. 11 10. 894 1,023. 75 7. 4710 1. 2459 1. 8204 2006 . 7535 1. 1340 7. 9723 1. 2563 116. 31 10. 906 954. 32 7. 3718 1. 2532 1. 8434 2007 . 8391 1. 0734 7. 6058 1. 3711 117. 76 10. 928 928. 97 6. 7550 1. 1999 2. 0020 2007: I. .7865 1. 1718 7. 7582 1. 3109 119. 33 11. 024 938. 98 7. 0089 1. 2330 1. 9548 II. .8316 1. 0983 7. 6784 1. 3484 120. 80 10. 878 928. 69 6. 8641 1. 2221 1. 9862 III. .8471 1. 0456 7. 5578 1. 3748 117. 74 10. 965 927. 27 6. 7402 1. 1986 2. 0213 IV. .8898 . 9811 7. 4336 1. 4482 113. 23 10. 849 921. 26 6. 4148 1. 1468 2. 0442 2008: I. .9058 1. 0039 7. 1590 1. 5007 105. 17 10. 803 956. 12 6. 2668 1. 0670 1. 9790 II. .9435 1. 0099 6. 9578 1. 5625 104. 62 10. 428 1,017. 02 5. 9862 1. 0316 1. 9712 III . 8879 1. 0411 6. 8375 1. 5030 107. 58 10. 328 1,064. 56 6. 3175 1. 0734 1. 8924 Table 3: Foreign exchange rates between 2000 and 2008 Source: Modified from GPO Access The strong dollar against the yen between 2001 and 2002, which was the beginning of the Bush administration, discouraged importation by Japan from the United States (Atkinson and Hutto). At the same time, citizens of the United States opted to import valuable items such as automobiles at the expense of the locally manufactured ones, thus putting the United States automobile industry in the quagmire in which it is today (OECD). The consequence of this is that major competitors such as the Japanese automakers have adversely affected local automakers such as Ford, and the whole industry has been earmarked for revival in the Economic Stimulus Package (OECD). Interest rates Government bond yields and interest rates generally declined between 2000 and 2008. For instance, the value of bills at auction declined steadily from $ 5. 85 to $1. 01 in 2003, but steadily rose from $1. 38 to $4. 73 in 2006 (GPO Access). The low price of bonds meant that banks were in a position to purchase more government bonds, thus diverting their attention from other financial service users such as borrowers (Crutsinger and Aversa). Ultimately, banks were forced to impose high interest rates on the loans they offered to the public, implying that most small business holders and individuals were crippled by a massive credit crunch (Crutsinger and Aversa). The ensuing credit crunch adversely affected the United States economy particularly towards the ends of the end of the Bush administration (Crutsinger and Aversa). Banks were most affected by the financial woes and this necessitated the Bush administration to consider taking ownership of various banks in a bid to protect them from collapsing (Crutsinger and Aversa). This move was however also ill planned, as it would result in unnecessary government expenditure, resulting into higher inflation in the United States (Hanke). In addition, there was no assurance that with the government’s acquisition of the banks, their performance would improve (Hanke). Conclusion. Even though the world witnessed a massive economic recession, the woes in the United States stemmed from the fact that the Bush administration spent excessively on the war in Iraq and Afghanistan, which generally weakened the performance of most local institutions. The underperformance in various sectors was shown in high rates of unemployment, high inflation, fluctuating levels of money supply and foreign exchange rates as well as high bank interest rates. The combined effect of the various phenomena led to a crisis in the entire United States economy, thus making the Bush administration one of the worst leaderships of the United States. References Atkinson Robert D and Julie Hutto 18October 2004. Bush vs. Clinton: An Economic Performance Index. 10 March 2009. http://www. ppionline. org/ppi_ci. cfm? knlgAreaID=107subsecID=295contentID=252964 Crutsinger, Martin and Aversa, Jeannine 8 October 2008. Bush administration mulls bank stakes. 10 March 2009. http://www. freep. com/article/20081008/BUSINESS07/81008120/1015/BUSINESS02/Bush+administration+mulls+bank+stakes Curl, Joseph. 23 Oct. 2008 McCain lambastes Bush years. 10 March 2009. http://www. washingtontimes. com/news/2008/oct/23/mccain-lambastes-bush-years/ GPO Access . Civilian unemployment rate. 10 March 2009. http://www. gpoaccess. gov/eop/2009/B42. xls GPO Access . Money stock and debt measures, 1965–2008. 10 March 2009 http://www. gpoaccess. gov/eop/2009/B69. xls GPO Access. Bond yields and interest rates, 1929–2008. 10 March 2009. http://www. gpoaccess. gov/eop/2009/B73. xls GPO Access. Changes in consumer price indexes for commodities and services, 1929–2007. 10 March 2009. http://www. gpoaccess. gov/eop/2009/B64. xls GPO Access. Foreign exchange rates, 1985–2008. 10 March 2009. http://www. gpoaccess. gov/eop/2009/B110. xls. Hanke, Steve H. September 24, 2008 The Bush Legacy: Deflation or Inflation? 10 March 2009. http://www. cato. org/pub_display. php? pub_id=9663 Irwin, Neil and Eggen, Dan. 12 Jan. 2009. The Washington Post. Economy Made Few Gains in Bush Years: Eight-Year Period Is Weakest in Decades. http://www. washingtonpost. com/wp-dyn/content/story/2009/01/12/ST2009011200359. html National Organization for Women. 27 August 2004. Bushs Economic Failure Weakens Middle Class, Deepens Poverty and Harms Women and Families. 10 March 2009. http://www. now. org/press/08-04/08-27. html OECD. Macroeconomic indicators. 10 March 2009. http://stats. oecd. org/mei/default. asp? lang=esubject=15country=USA Robinson, Woodward, Gellman. Timeline: Bushs Eight Years in Office. 10 March 2009. http://www. washingtonpost. com/wp-srv/politics/bush/legacy/timeline Shi, Leiyu and Stevens Gregory D. Vulnerable Populations in the United States. New York: John Wiley and Sons, 2004 The High Cost of Health Care. 25 November 2007. The New York Times. 10 March 2009. http://www. nytimes. com/2007/11/25/opinion/25sun1. html? _r=1 United states Department of Labor Bureau of Labor Statistics. 10 March 2009 http://www. bls. gov/.

Single Stage to Orbit (SSTO) Propulsion System

Single Stage to Orbit (SSTO) Propulsion System This paper discusses the relevant selection criteria for a single stage to orbit (SSTO) propulsion system and then reviews the characteristics of the typical engine types proposed for this role against these criteria. The engine types considered include Hydrogen/Oxygen (H2/O2) rockets, Scramjets, Turbojets, Turborockets and Liquid Air Cycle Engines. In the authors opinion none of the above engines are able to meet all the necessary criteria for an SSTO propulsion system simultaneously. However by selecting appropriate features from each it is possible to synthesise a new class of engines which are specifically optimised for the SSTO role. The resulting engines employ precooling of the airstream and a high internal pressure ratio to enable a relatively conventional high pressure rocket combustion chamber to be utilised in both airbreathing and rocket modes. This results in a significant mass saving with installation advantages which by careful design of the cycle thermodynamics enable s the full potential of airbreathing to be realised. The SABRE engine which powers the SKYLON launch vehicle is an example of one of these so called Precooled hybrid airbreathing rocket engines and the conceptual reasoning which leads to its main design parameters are described in the paper. Keywords: Reusable launchers, SABRE, SKYLON, SSTO 1.Introduction Several organisations world-wide are studying the technical and commercial feasibility of reusable SSTO launchers. This new class of vehicles appear to offer the tantalising prospect of greatly reduced recurring costs and increased reliability compared to existing expendable vehicles. However achieving this breakthrough is a difficult task since the attainment of orbital velocity in a re-entry capable single stage demands extraordinary propulsive performance. Most studies to date have focused on high pressure hydrogen/oxygen (H2/O2) rocket engines for the primary propulsion of such vehicles. However it is the authors opinion that despite recent advances in materials technology such an approach is not destined to succeed, due to the relatively low specific impulse of this type of propulsion. Airbreathing engines offer a possible route forward with their intrinsically higher specific impulse. However their low thrust/weight ratio, limited Mach number range and high dynamic pressure trajectory have in the past cancelled any theoretical advantage. By design review of the relevant characteristics of both rockets and airbreathing engines this paper sets out the rationale for the selection of deeply precooled hybrid airbreathing rocket engines for the main propulsion system of SSTO launchers as exemplified by the SKYLON vehicle [1]. 2. Propulsion Candidates This paper will only consider those engine types which would result in politically and environmentally acceptable vehicles. Therefore engines employing nuclear reactions (eg: onboard fission reactors or external nuclear pulse) and chemical engines with toxic exhausts (eg: fluorine/oxygen) will be excluded. The candidate engines can be split into two broad groups, namely pure rockets and engines with an airbreathing component. Since none of the airbreathers are capable of accelerating an SSTO vehicle all the way to orbital velocity, a practical vehicle will always have an onboard rocket engine to complete the ascent. Therefore the use of airbreathing has always been proposed within the context of improving the specific impulse of pure rocket propulsion during the initial lower Mach portion of the trajectory. Airbreathing engines have a much lower thrust/ weight ratio than rocket engines (à ¢Ã¢â‚¬ °Ã‹â€ 10%) which tends to offset the advantage of reduced fuel consumption. Therefore vehicles with airbreathing engines invariably have wings and employ a lifting trajectory in order to reduce the installed thrust requirement and hence the airbreathing engine mass penalty. The combination of wings and airbreathing engines then demands a low flat trajectory (compared to a ballistic rocket trajectory) in order to maximise the installed performance (i.e. (thrust-drag)/fuel flow). This high dynamic pressure trajectory gives rise to one of the drawbacks of an airbreathing approach since the airframe heating and loading are increased during the ascent which ultimately reflects in increased structure mass. However the absolute level of mass growth depends on the relative severity of the ascent as compared with reentry which in turn is mostly dependant on the type of airbreathing engine selected. An a dditional drawback to the low trajectory is increased drag losses particularly since the vehicle loiters longer in the lower atmosphere due to the lower acceleration, offset to some extent by the much reduced gravity loss during the rocket powered ascent. Importantly however, the addition of a set of wings brings more than just performance advantages to airbreathing vehicles. They also give considerably increased abort capability since a properly configured vehicle can remain in stable flight with up to half of its propulsion systems shutdown. Also during reentry the presence of wings reduces the ballistic coefficient thereby reducing the heating and hence thermal protection system mass, whilst simultaneously improving the vehicle lift/drag ratio permitting greater crossrange. The suitability of the following engines to the SSTO launcher role will be discussed since these are representative of the main types presently under study within various organisations world-wide: Liquid Hydrogen/Oxygen rockets Ramjets and Scramjets Turbojets/Turborockets and variants Liquid Air Cycle Engines (LACE) and Air Collection Engines (ACE) Precooled hybrid airbreathing rocket engines (RB545/SABRE) 3.Selection Criteria The selection of an optimum propulsion system involves an assessment of a number of interdependant factors which are listed below. The relative importance of these factors depends on the severity of the mission and the vehicle characteristics. Engine performance Useable Mach number and altitude range. Installed specific impulse. Installed thrust/weight. Performance sensitivity to component level efficiencies. Engine/Airframe integration Effect on airframe layout (Cg/Cp pitch trim structural efficiency). Effect of required engine trajectory (Q and heating) on airframe technology/materials. Technology level Materials/structures/aerothermodynamic and manufacturing technology. Development cost Engine scale and technology level. Complexity and power demand of ground test facilities. Necessity of an X plane research project to precede the main development program. 4.Hydrogen/Oxygen Rocket Engines Hydrogen/oxygen rocket engines achieve a very high thrust/weight ratio (60-80) but relatively low specific impulse (450-475 secs in vacuum) compared with conventional airbreathing engines. Due to the relatively large à ¢Ã‹â€ Ã¢â‚¬  V needed to reach low earth orbit (approx 9 km/s including gravity and drag losses) in relation to the engine exhaust velocity, SSTO rocket vehicles are characterised by very high mass ratios and low payload fractions. The H2/O2 propellant combination is invariably chosen for SSTO rockets due to its higher performance than other alternatives despite the structural penalties of employing a very low density cryogenic fuel. In order to maximise the specific impulse, high area ratio nozzles are required which inevitably leads to a high chamber pressure cycle in order to give a compact installation and reduce back pressure losses at low altitude. The need to minimise back pressure losses normally results in the selection of some form of altitude compensating nozzle since conventional bell nozzles have high divergence and overexpansion losses when running in a separated condition. The high thrust/weight and low specific impulse of H2/O2 rocket engines favours vertical takeoff wingless vehicles since the wing mass and drag penalty of a lifting trajectory results in a smaller payload than a steep ballistic climb out of the atmosphere. The ascent trajectory is therefore extremely benign (in terms of dynamic pressure and heating) with vehicle material selection determined by re-entry. Relative to airbreathing vehicles a pure rocket vehicle has a higher density (gross take off weight/volume) due to the reduced hydrogen consumption which has a favourable effect on the tankage and thermal protection system mass. In their favour rocket engines represent broadly known (current) technology, are ground testable in simple facilities, functional throughout the whole Mach number range and physically very compact resulting in good engine/airframe integration. Abort capability for an SSTO rocket vehicle would be achieved by arranging a high takeoff thrust/weight ratio (eg: 1.5) and a large number of engines (eg: 10) to permit shutdown of at least two whilst retaining overall vehicle control. From an operational standpoint SSTO rockets will be relatively noisy since the high takeoff mass and thrust/weight ratio results in an installed thrust level up to 10 times higher than a well designed airbreather. Reentry should be relatively straightforward providing the vehicle reenters base first with active cooling of the engine nozzles and the vehicle base. However the maximum lift/drag ratio in this attitude is relatively low (approx 0.25) limiting the maximum achievable crossrange to around 250 km. Having reached a low altitude some of the main engines would be restarted to control the subsonic descent before finally effecting a tailfirst landing on legs. Low crossrange is not a particular problem providing the vehicle operator has adequate time to wait for the orbital plane to cross the landing site. However in the case of a military or commercial operator this could pose a serious operational restriction and is consequently considered to be an undesirable characteristic for a new launch vehicle. In an attempt to increase the crossrange capability some designs attempt nosefirst re-entry of a blunt cone shaped vehicle or alternatively a blended wing/body configuration. This approach potentially increases the lift/drag ratio by reducing the fuselage wave drag and/or increasing the aerodynamic lift generation. However the drawback to this approach is that the nosefirst attitude is aerodynamically unstable since the aft mounted engine package pulls the empty center of gravity a considerable distance behind the hypersonic center of pressure. The resulting pitching moment is difficult to trim without adding nose ballast or large control surfaces projecting from the vehicle base. It is expected that the additional mass of these components is likely to erode the small payload capability of this engine/vehicle combination to the point where it is no longer feasible. Recent advances in materials technology (eg: fibre reinforced plastics and ceramics) have made a big impact on the feasibility of these vehicles. However the payload fraction is still very small at around 1-2% for an Equatorial low Earth orbit falling to as low as 0.25% for a Polar orbit. The low payload fraction is generally perceived to be the main disadvantage of this engine/vehicle combination and has historically prevented the development of such vehicles, since it is felt that a small degree of optimism in the preliminary mass estimates may be concealing the fact that the real payload fraction is negative. One possible route forward to increasing the average specific impulse of rocket vehicles is to employ the atmosphere for both oxidiser and reaction mass for part of the ascent. This is an old idea dating back to the 1950s and revitalised by the emergence of the BAe/Rolls Royce HOTOL project in the 1980s [2]. The following sections will review the main airbreathing engine candidates and trace the design background of precooled hybrid airbreathing rockets. 5.Ramjet and Scramjet Engines A ramjet engine is from a thermodynamic viewpoint a very simple device consisting of an intake, combustion and nozzle system in which the cycle pressure rise is achieved purely by ram compression. Consequently a separate propulsion system is needed to accelerate the vehicle to speeds at which the ramjet can takeover (Mach 1-2). A conventional hydrogen fuelled ramjet with a subsonic combustor is capable of operating up to around Mach 5-6 at which point the limiting effects of dissociation reduce the effective heat addition to the airflow resulting in a rapid loss in nett thrust. The idea behind the scramjet engine is to avoid the dissociation limit by only partially slowing the airstream through the intake system (thereby reducing the static temperature rise) and hence permitting greater useful heat addition in the now supersonic combustor. By this means scramjet engines offer the tantalising prospect of achieving a high specific impulse up to very high Mach numbers. The consequent de crease in the rocket powered à ¢Ã‹â€ Ã¢â‚¬  V would translate into a large saving in the mass of liquid oxygen required and hence possibly a reduction in launch mass. Although the scramjet is theoretically capable of generating positive nett thrust to a significant fraction of orbital velocity it is unworkable at low supersonic speeds. Therefore it is generally proposed that the internal geometry be reconfigured to function as a conventional ramjet to Mach 5 followed by transition to scramjet mode. A further reduction of the useful speed range of the scramjet results from consideration of the nett vehicle specific impulse ((thrust-drag)/fuel flow) in scramjet mode as compared with rocket mode. This tradeoff shows that it is more effective to shut the scramjet down at Mach 12-15 and continue the remainder of the ascent on pure rocket power. Therefore a scramjet powered launcher would have four main propulsion modes: a low speed accelerator mode to ramjet followed by scramjet and finally rocket mode. The proposed low speed propulsor is often a ducted ejector rocket system employing the scramjet injector struts as both ejector nozzles to entrain air at low speeds and later as the rocket combustion chambers for the final ascent. Whilst the scramjet engine is thermodynamically simple in conception, in engineering practice it is the most complex and technically demanding of all the engine concepts discussed in this paper. To make matters worse many studies including the recent ESA Winged Launcher Concept study have failed to show a positive payload for a scramjet powered SSTO since the fundamental propulsive characteristics of scramjets are poorly suited to the launcher role. The low specific thrust and high specific impulse of scramjets tends to favour a cruise vehicle application flying at fixed Mach number over long distances, especially since this would enable the elimination of most of the variable geometry. Scramjet engines have a relatively low specific thrust (nett thrust/airflow) due to the moderate combustor temperature rise and pressure ratio, and therefore a very large air mass flow is required to give adequate vehicle thrust/weight ratio. However at constant freestream dynamic head the captured air mass flow reduces for a given intake area as speed rises above Mach 1. Consequently the entire vehicle frontal area is needed to serve as an intake at scramjet speeds and similarly the exhaust flow has to be re-expanded back into the original streamtube in order to achieve a reasonable exhaust velocity. However employing the vehicle forebody and aftbody as part of the propulsion system has many disadvantages: The forebody boundary layer (up to 40% of the intake flow) must be carried through the entire shock system with consequent likelihood of upsetting the intake flow stability. The conventional solution of bleeding the boundary layer off would be unacceptable due to the prohibitive momentum drag penalty. The vehicle undersurface must be flat in order to provide a reasonably uniform flowfield for the engine installation. The flattened vehicle cross section is poorly suited to pressurised tankage and has a higher surface area/volume than a circular cross section with knock-on penalties in aeroshell, insulation and structure mass. Since the engine and airframe are physically inseparable little freedom is available to the designer to control the vehicle pitch balance. The single sided intake and nozzle systems positioned underneath the vehicle generate both lift and pitching moments. Since it is necessary to optimise the intake and nozzle system geometry to maximise the engine performance it is extremely unlikely that the vehicle will be pitch balanced over the entire Mach number range. Further it is not clear whether adequate CG movement to trim the vehicle could be achieved by active propellant transfer. Clustering the engines into a compact package underneath the vehicle results in a highly interdependant flowfield. An unexpected failure in one engine with a consequent loss of internal flow is likely to unstart the entire engine installation precipitating a violent change in vehicle pitching moment. In order to focus the intake shock system and generate the correct duct flow areas over the whole Mach range, variable geometry intake/combustor and nozzle surfaces are required. The large variation in flow passage shape forces the adoption of a rectangular engine cross section with flat moving ramps thereby incurring a severe penalty in the pressure vessel mass. Also to maximise the installed engine performance requires a high dynamic pressure trajectory which in combination with the high Mach number imposes severe heating rates on the airframe. Active cooling of significant portions of the airframe will be necessary with further penalties in mass and complexity. Further drawbacks to the scramjet concept are evident in many areas. The nett thrust of a scramjet engine is very sensitive to the intake, combustion and nozzle efficiencies due to the exceptionally poor work ratio of the cycle. Since the exhaust velocity is only slightly greater than the incoming freestream velocity a small reduction in pressure recovery or combustion efficiency is likely to convert a small nett thrust into a small nett drag. This situation might be tolerable if the theoretical methods (CFD codes) and engineering knowledge were on a very solid footing with ample correlation of theory with experiment. However the reality is that the component efficiencies are dependant on the detailed physics of poorly understood areas like flow turbulence, shock wave/boundary layer interactions and boundary layer transition. To exacerbate this deficiency in the underlying physics existing ground test facilities are unable to replicate the flowfield at physically representative sizes , forcing the adoption of expensive flight research vehicles to acquire the necessary data. Scramjet development could only proceed after a lengthy technology program and even then would probably be a risky and expensive project. In 1993 Reaction Engines estimated that a 130 tonne scramjet vehicle development program would cost $25B (at fixed prices) assuming that the program proceeded according to plan. This program would have included two X planes, one devoted to the subsonic handling and low supersonic regime and the other an air dropped scramjet research vehicle to explore the Mach 5-15 regime. 6.Turbojets, Turborockets and Variants In this section are grouped those engines that employ turbocompressors to compress the airflow but without the aid of precoolers. The advantage of cycles that employ onboard work transfer to the airflow is that they are capable of operation from sea level static conditions. This has important performance advantages over engines employing solely ram compression and additionally enables a cheaper development program since the mechanical reliability can be acquired in relatively inexpensive open air ground test facilities. 6.1 Turbojets Turbojets (Fig. 1) exhibit a very rapid thrust decay above about Mach 3 due to the effects of the rising compressor inlet temperature forcing a reduction in both flow and pressure ratio. Compressors must be operated within a stable part of their characteristic bounded by the surge and choke limits. In addition structural considerations impose an upper outlet temperature and spool speed limit. As inlet temperature rises (whilst operating at constant Wà ¢Ã‹â€ Ã… ¡T/P and N/à ¢Ã‹â€ Ã… ¡T) the spool speed and/or outlet temperature limit is rapidly approached. Either way it is necessary to throttle the engine by moving down the running line, in the process reducing both flow and pressure ratio. The consequent reduction in nozzle pressure ratio and mass flow results in a rapid loss in nett thrust. However at Mach 3 the vehicle has received an insufficient boost to make up for the mass penalty of the airbreathing engine. Therefore all these cycles tend to be proposed in conjunction with a subsonic combustion ramjet mode to higher Mach numbers. The turbojet would be isolated from the hot airflow in ramjet mode by blocker doors which allow the airstream to flow around the core engine with small pressure loss. The ramjet mode provides reasonable specific thrust to around Mach 6-7 at which point transition to rocket propulsion is effected. Despite the ramjet extension to the Mach number range the performance of these systems is poor due mainly to their low thrust/weight ratio. An uninstalled turbojet has a thrust/weight ratio of around 10. However this falls to 5 or less when the intake and nozzle systems are added which compares badly with a H2/O2 rocket of 60+. 6.2 Turborocket The turborocket (Fig. 2) cycles represent an attempt to improve on the low thrust/weight of the turbojet and to increase the useful Mach number range. The pure turborocket consists of a low pressure ratio fan driven by an entirely separate turbine employing H2/O2 combustion products. Due to the separate turbine working fluid the matching problems of the turbojet are eased since the compressor can in principle be operated anywhere on its characteristic. By manufacturing the compressor components in a suitable high temperature material (such as reinforced ceramic) it is possible to eliminate the ramjet bypass duct and operate the engine to Mach 5-6 whilst staying within outlet temperature and spool speed limits. In practice this involves operating at reduced nondimensional speed N/à ¢Ã‹â€ Ã… ¡T and hence pressure ratio. Consequently to avoid choking the compressor outlet guide vanes a low pressure ratio compressor is selected (often only 2 stages) which permits operation over a wider flow range. The turborocket is considerably lighter than a turbojet. However the low cycle pressure ratio reduces the specific thrust at low Mach numbers and in conjunction with the preburner liquid oxygen flow results in a poor specific impulse compared to the turbojet. 6.3 Expander Cycle Turborocket This cycle is a variant of the turborocket whereby the turbine working fluid is replaced by high pressure regeneratively heated hydrogen warmed in a heat exchanger located in the exhaust duct (Fig. 3). Due to heat exchanger metal temperature limitations the combustion process is normally split into two stages (upstream and downstream of the ma- LHLH LOx/LH2 Fig. 1 Turbo-ramjet Engine (with integrated rocket engine). LOx/LH2LH2 LOx/LH2 Fig. 2 Turborocket. LH2LOx/LH2 Fig. 3 Turbo-expander engine. trix) and the turbine entry temperature is quite low at around 950K. This variant exhibits a moderate improvement in specific impulse compared with the pure turborocket due to the elimination of the liquid oxygen flow. However this is achieved at the expense of additional pressure loss in the air ducting and the mass penalty of the heat exchanger. Unfortunately none of the above engines exhibit any performance improvement over a pure rocket approach to the SSTO launcher problem, despite the wide variations in core engine cycle and machinery. This is for the simple reason that the core engine masses are swamped by the much larger masses of the intake and nozzle systems which tend to outweigh the advantage of increased specific impulse. Due to the relatively low pressure ratio ramjet modes of these engines, it is essential to provide an efficient high pressure recovery variable geometry intake and a variable geometry exhaust nozzle. The need for high pressure recovery forces the adoption of 2 dimensional geometry for the intake system due to the requirement to focus multiple oblique shockwaves over a wide mach number range. This results in a very serious mass penalty due to the inefficient pressure vessel cross section and the physically large and complicated moving ramp assembly with its high actuation loads. Similarly the exhaust nozzle geometry must be capable of a wide area ratio variation in order to cope with the widely differing flow conditions (Wà ¢Ã‹â€ Ã… ¡T/P and pressure ratio) between transonic and high Mach number flight. A further complication emerges due to the requirement to integrate the rocket engine needed for the later ascent into the airbreathing engine nozzle. This avoids the prohibitive base drag penalty that would result from a separate dead nozzle system as the vehicle attempted to accelerate through transonic. 7. Liquid Air Cycle Engines (LACE) and Air Collection Engines (ACE) Liquid Air Cycle Engines were first proposed by Marquardt in the early 1960s. The simple LACE engine exploits the low temperature and high specific heat of liquid hydrogen in order to liquify the captured airstream in a specially designed condenser (Fig. 4). Following liquifaction the air is relatively easily pumped up to such high pressures that it can be fed into a conventional rocket combustion chamber. The main advantage of this approach is that the airbreathing and rocket propulsion systems can be combined with only a single nozzle required for both modes. This results in a mass saving and a compact installation with efficient base area utilisation. Also the engine is in principle capable of operation from sea level static conditions up to perhaps Mach 6-7. LH2 LO2 Liquid Air Turbopump Fig. 4 Liquid Air Cycle Engine (LACE). The main disadvantage of the LACE engine however is that the fuel consumption is very high (compared to other airbreathing engines) with a specific impulse of only about 800 secs. Condensing the airflow necessitates the removal of the latent heat of vaporisation under isothermal conditions. However the hydrogen coolant is in a supercritical state following compression in the turbopump and absorbs the heat load with an accompanying increase in temperature. Consequently a temperature pinch point occurs within the condenser at around 80K and can only be controlled by increasing the hydrogen flow to several times stoichiometric. The air pressure within the condenser affects the latent heat of vaporisation and the liquifaction temperature and consequently has a strong effect on the fuel/air ratio. However at sea level static conditions of around 1 bar the minimum fuel/air ratio required is about 0.35 (ie: 12 times greater than the stoichiometric ratio of 0.029) assuming that the hydrogen had been compressed to 200 bar. Increasing the air pressure or reducing the hydrogen pump delivery pressure (and temperature) could reduce the fuel/ air ratio to perhaps 0.2 but nevertheless the fuel flow remains very high. At high Mach numbers the fuel flow may need to be increased further, due to heat exchanger metal temperature limitations (exacerbated by hydrogen embrittlement limiting the choice of tube materials). To reduce the fuel flow it is sometimes proposed to employ slush hydrogen and recirculate a portion of the coolant flow back into the tankage. However the handling of slush hydrogen poses difficult technical and operational problems. From a technology standpoint the main challenges of the simple LACE engine are the need to prevent clogging of the condenser by frozen carbon dioxide, argon and water vapour. Also the ability of the condenser to cope with a changing g vector and of designing a scavenge pump to operate with a very low NPSH inlet. Nevertheless performance studies of SSTOs equipped with LACE engines have shown no performance gains due to the inadequate specific impulse in airbreathing mode despite the reasonable thrust/weight ratio and Mach number capability. The Air Collection Engine (ACE) is a more complex variant of the LACE engine in which a liquid oxygen separator is incorporated after the air liquifier. The intention is to takeoff with the main liquid oxygen tanks empty and fill them during the airbreathing ascent thereby possibly reducing the undercarriage mass and installed thrust level. The ACE principal is often proposed for parallel operation with a ramjet main propulsion system. In this variant the hydrogen fuel flow would condense a quantity of air from which the oxygen would be separated before entering the ramjet combustion chamber at a near stoichiometric mixture ratio. The liquid nitrogen from the separator could perform various cooling duties before being fed back into the ramjet airflow to recover the momentum drag. The oxygen separator would be a complex and heavy item since the physical properties of liquid oxygen and nitrogen are very similar. However setting aside the engineering details, the basic thermodynamics of the ACE principal are wholly unsuited to an SSTO launcher. Since a fuel/air mixture ratio of approximately 0.2 is needed to liquify the air and since oxygen is 23.1% of the airflow it is apparent that a roughly equal mass of hydrogen is required to liquify a given mass of oxygen. Therefore there is no saving in the takeoff propellant loading and in reality a severe structure mass penalty due to the increased fuselage volume needed to contain the low density liquid hydrogen. 8. Precooled Hybrid Airbreathing Rocket Engines This last class of engines is specifically formulated for the SSTO propulsion role and combines some of the best features of the previous types whilst simultaneously overcoming their faults. The first engine of this type was the RB545 powerpla

Father and Son by Bernard MacLaverty Essay example -- English Literatu

Father and Son By Bernard MacLaverty Father and Son is a story about the relationship between a widower father and his teenage son. â€Å"Father and Son† â€Å"Father and Son† is a story about the relationship between a widower father and his teenage son. They live in Belfast, in a neighbourhood with a lot of violence. At night they can hear the sound of ambulances criss-crosses the dark. Both the son and the father are scared to sleep at night, but the son will not admit his fears to his father. The father is concerned for his son’s life. They used to have a good relationship. They went fishing, talked and laughed, and the father could put his arms around his son. But two years before this story takes place the son went to London and got mixed up with drug addicts. His father had t...

Critical Analysis of Shakespeares Hamlet Essay -- Shakespeare Hamlet

Critical Analysis of Shakespeare's Hamlet What is mans' purpose in life? Is there a purpose? If there isn't, then is it wise to end it, despite the fact that there might be nothing better? In Hamlet, by William Shakespeare, Hamlet struggles with these and other issues. He states that the question of life is "To be, or not to be...?" Is existence really worth the troubles of life? In this monologue, Hamlet is wondering what is his purpose. He asserts that the only reason people endure their horrible lives is the uncertainty of what lies after death. "Who would fardels bear, to grunt and sweat under a weary life, but that the dread of something after death..." Is it noble to suffer, and is life worth all its misery? Hamlet must question himself to discover the answers. At the point in Hamlet when this famous soliloquy takes place, Hamlet has many reasons to be questioning his existence. Hamlet is visited by the ghost of his late father, who explains that he was murdered by Hamlet's uncle, who is Hamlet's mother's new husband. His father cannot rest until Hamlet has gotten revenge. Hamlet's father has just been murdered, his friends are sent to spy on him, his lover is forbidden to see him, and Hamlet feels that his life is pointless and miserable. "The pangs of disprized love, the law's delay, the insolence of office, and the spurns that patient merit of the unworthy takes..." These are the miseries that Hamlet must endure. This is why he makes this s...

John Steinbeck’s Of Mice and Men Essay -- Literary Criticism

Abstract: When Of Mice and Men is read between the lines, it is easy to discover Steinbeck’s objective. Men are destined to be alone, despite their hardships, despite their connections. In order to establish these ideas Steinbeck develops his characters, the relationships they share, and their interactions. Of Mice and Men is considered a classic novel for a reason. Steinbeck does an excellent job of establishing powerful themes while using unparalleled writing techniques. Steinbeck was able to take his upbringing as the common man and turn it into inspiration for novels. At the mention of the name, â€Å"John Steinbeck,† many associations can be made to the classic works produced by the man. However, he did not begin his life as a successful writer. Growing up in Salinas, California, Steinbeck lived the life of a common man, working to survive in the Land of Promise. He began to develop a taste for writing; however, he studied marine biology while he attended Stanford University. Without graduating with a degree, Steinbeck began working as a laborer and reporter for the American in New York City. After realizing that he was not meeting his goals, or at least coming close to them, Steinbeck moved back to California where he continued to work in various odd jobs while he pursuing his career as a professional writer. In the early 1930’s Steinbeck met Edwards Ricketts, a marine biologist who later became a major influence on his works. Steinbeck enjoyed listening to Ricketts’s views on the interdependence of life. Steinbeck became i ntrigued by these concepts and began applying these themes to all of his developing works. His first two publications were undeveloped and deemed failures, but with the production of Of Mice and Men, Steinbeck was introduced to the world of literary acceptance.   Ã‚  Ã‚  Ã‚  Ã‚  Of Mice and Men is considered an American classic that is taught by almost every high school and featured on most â€Å"must read† booklists. The reason for such appraise can be found in the themes lying within the book. Steinbeck establishes the concepts of man’s destiny by developing his characters, the relationships they share, and their interactions and using them to reinforce underlying themes. The relationship between Lennie and George opens many types of interpretation for critics of the novel. Some wonder whether George cares for Lennie and a person, or maybe Geo... ...e alone, despite their hardships, despite their connections. In order to establish these ideas Steinbeck develops his characters, the relationships they share, and their interactions. Of Mice and Men is considered a classic novel for a reason. Steinbeck does an excellent job of establishing powerful themes while using unparalleled writing techniques. Steinbeck was able to take his upbringing as the common man and turn it into inspiration for novels. WORKS CITED Attell, Kevin. Novels for Students: Man and Animal. Stanford University Press: Stanford California, 2004. Folsom, James K.. Critical Survey of Short Fiction. Salem Press: Englewood Cliffs, NJ, 1981. French, Warren. Reference Guide to American Literature. Harmony Books: New York, 1987 Goldhurst, William. Of Mice and Men: John Steinbeck’s Parable of the Curse of Cain. Western American Literature: New York, 1971. Levant, Howard. The Novels of John Steinbeck: A Critical Study. Columbia, Mo.: University of Missouri Press, 1974. Lisca, Peter. John Steinbeck, Nature and Myth. New York: Thomas Y. Crowell, 1978. Paul, Louis. Contemporary Literary Criticism. William Beneton: Detroit, Michigan, 1982.

The Truth Should Be Told, Whatever the Cost

Anisa Kornegay P. 7 Discursive Essay #2 â€Å"The truth should always be told, whatever the cost. † It’s clear that in life you run into lies and you run into the truth but when is it morally ethical to not tell the truth, and when is it okay to lie? The truth has caused a plethora of controversy worldwide instances where it is concealed and when it is exposed. Subsequently the truth seems to always be told, when it is and even when it is not needed. In the medical realm the truth should be told, although in the past it has been covered up.Medical experiments frequently avoid telling the truth. For example in the Tuskegee Syphilis experiment in Alabama, 399 black men were told they were being treated for â€Å"bad blood† when in actuality they were in the late stages of Syphilis. By the end of the experiment 128 men died from syphilis, but if the doctors informed the subjects of the infection many lives could have been spared. Just like the Tuskegee Syphilis expe riment, in other occasions the truth should be told in the medical world.For instance, in Crisis Pregnancy Centers all over the United States also known as CPC’s pregnant women are told claims by the uneducated staff members to have them refrain from having an abortion because they believe abortion is wrong. One woman, Katy Stag went to three CPC’s in Ohio that told her â€Å"There is more risk in having an abortion than carrying a baby full term† which is false; studies show that a woman who carries their baby full term are fourteen times more likely to die. This exemplifies that in the medical world that lies should not be told.Therefore, the truth should be told, especially in regard to anything medical. Along with the medical world, the truth should be told in regard to high government officials. Presidents should always tell the truth, one lie can lead a president to impeachment. For example, in 1974 President Richard Nixon of the United States was on the v erge of impeachment for obstructing justice. President Nixon hired men to illegally gain information on his opposing party. Nixon denied all involvement, even when evidence surfaced linking him to the crime.Just like President Richard Nixon Watergate scandal, other presidents have withheld the truth. For instance, Bill Clinton was impeached on the charge of obstruction of justice in 1998. The charges came from his sexual affair with his 22 year old intern at the time Monica Lewinsky, DNA evidence proved there was a sexual relationship even though President Clinton denied the allegations. His lie had him impeached. Evidently, the president of a country is one looked up to and should tell the truth. It doesn’t make a significance difference in citizens’ lives if an athlete lies.If the action of an athlete does not affect the lives of others the truth should not be of significance. For example in 1988, Gregory Louganis a diver in the Olympic Games, went to dive and hit hi s head on the platform, later the truth came out that he had HIV. AIDS expert Anthony Fauci, MD, assured everyone that Greg Louganis did not place any risk to the health of the other athletes on his decision to disclose his HIV diagnosis. Just like with Gregory Louganis other athletes have had the truth revealed when it was unnecessary.For instance, Marion Jones, 3 time Olympic gold medalist track star, plead guilty in 2004 to lying to federal agents about taking banned steroids, but it was proved her usage was not during any of the times she won her gold medals, so it did not change the Olympic results and she keeps her medals. Without the truth, nothing really would have changed. Subsequently, the lives of citizens are not affected when an athlete has lied about their situation leaving the truth better off not exposed. The effects of war or danger are a reason to withhold the truth. In warfare the truth is permitted to be disguised by lies to be loyal to your own country.For examp le, in the Cold war between NATO and the communist sect, people on each side of the war were prepared to betray their countries secrets, but would sacrifice themselves to protect their country. You do what you have to win the war for your country by any means, and if those means involve lying then so be it. A similar principle was present in the 1939; 45 year war giving out â€Å"disinformation† was a key role in the war. Hitler’s forces in France were deceived to think that the war was at the Pas de Calais rather than along the Normandy coast, England even created a fake army.This morale shows they did what they could to win the war. Undoubtedly, the danger of war outweigh telling the truth, thus it would be acceptable to lie. It is apparent that the truth has certain circumstances when it is to be exposed and when it is to be concealed. In medicine and at the head of government it seems the truth is to be told constantly. On the other hand, in sports and in warfare t he truth is not as needed. No matter what, one thing is certain, we as humans must learn the instances of when the truth should be told regardless of the consequences.