4.1 Earth Launch Vehicle

4.1 Earth Launch Vehicle

Stephen S. G. Hanna

Nomenclature

q = Dynamic Pressure

V = Velocity

CR = Coverage by a rescue system

PE = Relative Probability of a Single Emergency

QER = Quality Factor

TER = Time Factor

r = Density

4.1.1 Introduction for Earth Launch Vehicle

The mission directives focus is on placing astronauts on the Martian surface while minimizing cost and insuring a crew survival rate greater than 95% and a mission success rate greater than 80%. The mission is a "stepping stone" for future missions as a result a versatile system was chosen for the ELV. The vehicle required for Earth-to-Orbit is fundamentally guided by these criteria as well as the mass of the largest payload projected for the Martian surface.

4.1.2 Launch Systems

The hypothetical design mass assumed for vehicles sent to Mars is 65 tonnes for a crew habitat sized for four people that is transferred on a high-energy orbit based on historical data.1 The 65 tonne spacecraft does not include the propulsion system needed for sending the spacecraft on a trans Mars injection path. Thus the capability of the selected launch vehicle is approximately 200 tonnes to LEO.

A 200-ton-class launch vehicle raises many development cost issues. Consideration is given to the option of launching multiple parts of a space vehicle to LEO. Therefore, a smaller launcher is used and assembly of the parts in orbit prior to launching to Mars would occur. The smaller launch vehicle, 110 to 120 tonnes, would have the advantage of more modest development costs and is within the capability of existing technologies.

The smaller launch vehicle though appealing introduces several potential difficulties to the mission. The simplest, most desirable implementation of the smaller launch vehicle is to simply dock the two elements in Earth orbit and immediately depart for Mars. Avoiding boiloff loss of cryogenic propellants in the departure stages but forcing all elements launched from Earth to occur in a quick succession. Placing a large pressure on a single launch facility and its ground operations crews or the close coordination of two or more launch facilities. Assembling the Mars vehicles in orbit and loading them with propellants from an orbiting depot just prior to departure may alleviate the strain on the launch facilities. The best Earth orbit for a Mars mission is different for each launch opportunity in addition to the cost incurred to make a fueling depot in space.

Earth Launch Vehicle [ELV] options

For our mission; however, the larger, 200 tonne class launch vehicle is picked to help reduce complexity therefore improving reliability and success of the mission. Constructing a single launch vehicle capable of placing 200 tonnes into LEO can employ many different technologies. The launch vehicle concepts in Table 4.1.1 used various combinations of past, present, and future U.S. expendable launch vehicle technology and existing launch vehicle technology from Russia.1,3

Table 4.1.1 Launch Vehicle Options
Payload Mass (tonnes) / Estimated Cost
(In 2000 Dollars) / Key Technology
179 / $3 billion / Modified Energia core with eight Zenit-type strap-on boosters. Proton 4th stage for Retro Correction. All production facilities exist and can be licensed abroad.
209 / $8 billion / Space Shuttle C Derivative. New core stage based on Space Transportation System (STS) external tank and SSME. Seven new strap-on boosters each use a single RD-170 engine. New upper stage using a single SSME.
226 / $10 billion / New core stage based on STS external tank and four of the new Space Transportation Main Engines. Four strap-on boosters each with a derivative of the F-1 engine used on the first stage of the Saturn V. New upper stage using a single SSME.
289 / $12 billion / New vehicle using technology derived from the Saturn V launch vehicle. Boosters and first stage use a derivative of the F-1 engine, and the second stage uses a derivative of the J-2 engine. New launch facilities and Vehicle Assembly Building (VAB).

We decided that the Energia based vehicle for the ELV is a cost effective choice for our potential payload. The Energia is a modular heavy-lift launch vehicle capable of launching a variety of payloads in tandem with the NTR (Nuclear Thermal Rocket) as second stage making an exceedingly viable solution.

4.1.3 ELV- Energia

The ELV consists of a central core with four liquid hydrogen/ liquid oxygen engines and eight strap-on-boosters, each with a single four-chamber liquid oxygen/ liquid kerosene engine. Four RD-0120 engines comprise the main core of the Energia launch vehicle. RD-0120 rocket engine is the Russian equivalent to the US SSME (Space Shuttle Main Engine). The eight strap-on-boosters are composed of RD-170 engines. The RD-170 is a high altitude engine in the Zenit second stage. The RD-170 engine consists of 4 chambers, 1 turbo-pump and 2 gas generators; RD-170 uses a one-plane gimballing. The RD- 170 is the first Russian designed engine to be constructed and test fired in the United States. More detailed information is found in the Appendix 4.1 Matlab ‘rocket.m’ code.

The Energia engines are existing engines with heritage. A benefit of the engines is the capability to license production in the US. Another modification involves changing the vehicle from one using a side mounted payload container to an in- line arrangement with strap-on boosters surrounding the core. The core stage fairing is assumed to be a modification of the existing Energia stage using the US Space Shuttle External (SSE) tank technology and facilities. The capabilities of current facilities and the dimensions of both the SSE tank and the Energia agree. See Fig. 4.1.4 for a schematic of the ELV.

4.1.5 ELV Performance

All core and strap-on-boosters are ignited at liftoff; engine throttling is used to manage axial acceleration and dynamic pressure. The strap-on-booster and core employ the gimbaling capabilities of the engines, plus or minus 5 or 8.6 degrees respectively, to maintain attitude control. Following propellant depletion the core continues to burn until the desired shutdown state is reached. Further information is available in Fig. 4.1.2 and Table 4.1.2; all numbers are verified with historical sources and a Matlab ‘rocket.m’ code in the Appendix 4.1.

Table 4.1.2 ELV Flight Path and Performance
Event / Time (min:sec) / Altitude (km) / Down Range Distance / Inertial Velocity / Notes
1. Liftoff / 00:00 / 0 / N/A / N/A / Kennedy Space Center
2. Booster Staging / 2:20 / 80 / 400 (Atlantic Ocean) / 1.8 km/sec / Reusable Up to Ten times
3. Core Separation / 6:30 / 110 / 19200 (Indian Ocean) / 658,368 km/sec / Potentially
Reusable

The Energia is unique among launch vehicles in it does not deliver its payload directly to orbit. In order to facilitate a predictable and safe reentry of the core, the core engines are shutdown, and the core is jettisoned, before orbital velocity is reached. The final velocity increment required to achieve orbit is provided by the RCS (Retro- Correction Stage) based on the Proton 4th Stage. The attitude control of the RCS is maintained by an APE (auxiliary propulsion engine), which contains two sets of small fixed thrusters for three axes control.

4.1.6 Launch Vehicle Safety

The lives and safety of the astronauts should always have top priority in all manned space flight activities. However a 100% safety can never be reached.

As inherent risk is undertaken by the astronauts due to the to the nature of the technical devices such as the launch and space vehicles, whose reliability is never perfect, due to influences of the environment, like wind, low temperatures, and through human error. Thus an inherent risk remains in space flight through all efforts to reach a maximum of technical reliability. About 89% of all severe emergencies of space missions occur during launch.6 The risk of this relatively short phase is about as high as living and working on the space station for one year. 5

The installation of safety and rescue systems such as ejection systems or rescue capsules always leads to additional weight and causes a reduction in a payload capability. Due to relatively low launching rates it is hard to obtain exact safety data of manned space vehicles and launchers. Therefore the relative safety gains of a various vehicles are compared and then factored against the vehicles’ existing history. We accomplished this by investigating properties such as mission capabilities, weight and operational aspects. The main criterion is maximum safety without the dramatic reduction of payload to LEO.

4.1.7 Launch Vehicle Reliability

The ELV maintains some heritage as it is based on the Russian Energia. If the reliability of the launch vehicle is broken up by individual components; it makes for a better estimate to evaluate the vehicles reliability because there it gives more history to individual components.

The system is broken up into three major components main core, strap boosters, and retro- correction stage (RCS). The eight strap-on-boosters, similar to Russian Zenit first stage booster, have a 96% success rate found on twenty-three launches and one failure of the first stage.3 The Energia system is built on redundancy so a failure of one engine is acceptable; therefore, giving the boosters an 87.5% (1 out of 8) reliability. 3

The main core is comprised of four engines that in the current configuration do not carry any heritage other than two previous Energia launches. The engines themselves are human space rated; therefore, qualify for the launch of the astronauts. According to thrust calculations needed for payload insertion only three engines are needed for LEO insertion; therefore, giving the main core 75% reliability.

The overall reliability of the ELV is 87.5% for a successful mission based on booster performance with a 96% success rate based on booster heritage. Therefore to improve survivability rate at launch, a zero altitude abort is needed.

Rescue systems/ Zero Abort

Rescue systems are defined as a set of additional, often multiple redundant systems, and installations like life support, power supply, communication, guidance and control, caution, warning systems, and pressure suits. Zero altitude abort is the ability to safely retrieve the crew while the launch vehicle is on the pad (launch site) due to an emergency.8 We investigated various systems including pressurized suits, open ejection seats, encapsulated ejection seats, rescue cabin, and tower jets.

Pressurized Suits

During launch and re- entry astronauts will be required to wear pressurized suits to protect against loss of pressure up to an altitude of approximately 40 km. These suits also protect against extreme temperatures and dynamic pressure (Eq. 4.1.1) in case of an abort due to emergencies. They include autonomous oxygen and survival kits for each crewmember.

q = ½*r * V2 (4.1.1)

The suits are necessary to help in redundancy against leaks to the pressurized vessel as historically shown in the Soyuz 11 catastrophe.5 As the expense for personal safety is great and the weight of a pressurized suit is 10Kg per person; the choice of enlisting pressurized suits is practical.

Open ejection seats

Open ejection seats are self-contained by including all propulsive devices, autonomous oxygen, and parachutes (drone chute and main chute) and have been proven at varied speeds and altitudes. Ejection seats are viable if seats remain within the range operation. Though ejection seats are theoretically not viable higher than 40km because of the inefficiencies of the pressure suits. Ejection seats have been used at the heights of 90 km with survival of the pilots An added redundancy is the use of pressure suits along with ejection seats as parachutes are equipped with all pressure suits. Max dynamic pressure felt by the launch vehicle is 1400l b_ft2.6 Beyond this point, the pressure suits are not sufficient to protect the astronauts against dynamic pressure and do not provide adequate atmospheric conditions.

Ejection seats weigh in the range of 98 kg (Martin – Baker) to 225Kg (Burn) including pressure suit weights.9 Using the highest rated ejection seat for a crew of four the total weight of the system is 816 kg (204kg each*4 crew = 816 kg total). Further detail can be found in Appendix 4.1.2. As ejection seats replace normal launch seats, the seats increase HAB (Habitation module) weight by less than their own weight. The installation, however, implies changes of the vehicle construction, especially jettisonable panels. Therefore we assume the overall additional mass of installed ejection seats to be equal to the above –mentioned values. Almost fifty years experience with ejection seats in military aviation have led to the high level of efficiency and reliability of these systems, but they have never been used in space flight emergencies.

Encapsulated ejection seats and rescue cabin

There is minimal heritage to encapsulated ejections for aviation and none for space flight. Encapsulated ejection seats have the added weight of open ejection seats; the mass of the capsule devices and damping systems for landing is about 140 Kg in addition to the ejection seat.6 The added safety of an encapsulated ejection seat is minimal to the mass payload lost. Rescue cabins include two options ejecting the crew section or jettisoning the entire vehicle both options are impractical. The complexity of jettisoning the crew is not a justifiable due to complexity and cost. Jettisoning the entire vehicle is made difficult by the blast zone of the rocket of a minimum 1000 meters based on Mass propellant.10

Rescue Cabin Ejection Pod. The weight of implementation is mass of an ejection pod (not including propulsion system) 1400 kg + ejection seats 816 kg + pressurized suits 40 kg à 996kg. The system is not viable based on our spacecraft design. The HAB is built not only as re-entry vehicle, a spacecraft, but a habitat thus leading to large volume and weight constraints. The loss of volume for an ejection pod is high though payload weight loss is <20% of total payload weight.7 Ejection pod is not practical due to costly development costs. Because the pod must have a fly by wire system implemented and must also clear the blast zone. An ejection pod can cost up too one and half billion dollars (Based on Hermes craft cost of development). 8

Tower Jets L. For a tower jet to be safe it must reach a safety height of 1 km, due to the blast character of the launch vehicle is 1 km based on mass propellant and type.10 An ISP of a typical liquid engine is used for calculations though a solid rocket motor is the most practical for a tower jet. A burn time of 6 seconds is historically used for tower jets with a max acceleration of 12g’s for astronaut’s health.5 A structural mass of 10% of the tower jet for the attaching of the tower jet to the spacecraft is assumed. A mass of the spacecraft of 75 tonnes was used to calculate the tower jet weight. No drag or gravity is considered for preliminary analysis.